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.

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.

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.

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.

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.

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.

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.

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.

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.

Numerical simulation of turbulent flow and heat transfer though sinusoidal ducts

NASA Astrophysics Data System (ADS)

Abroshan, Hamid

2018-02-01

Turbulent forced convection heat transfer in corrugated plate surfaces was studied by means of CFD. Flow through corrugated plates, which are sets of sinusoidal ducts, was analyzed for different inlet flow angles (0° to 50°), aspect ratios (0.1 to 10), Reynolds numbers (2000 to 40,000) and Prantdel numbers (0.7 to 5). Heat transfer is affected significantly by variation of aspect ratio. A maximum heat transfer coefficient is observed at a particular aspect ratio although the aspect ratio has a minor effect on friction factor. Enlarging inlet flow angle also leads to a higher heat transfer coefficient and pressure loss in aspect ratios close to unity. Dependency of Nusselt and friction factor on the angle and aspect ratio was interpreted by means of appearance of secondary motions and coexistence of laminar and turbulent flow in a cross section. Comparing the results with experimental data shows a maximum 12.8% difference. By evaluating the results, some correlations were proposed to calculate Nusselt number and friction factor for entrance and fully developed regions. A corrugated plate with an aspect ratio equal to 1.125 and an inlet flow angle equal to 50° gives the best heat transfer and pressure drop characteristics.

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.

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.

Modal propagation angles in a cylindrical duct with flow and their relation to sound radiation

NASA Technical Reports Server (NTRS)

Rice, E. J.; Heidmann, M. F.; Sofrin, T. G.

1979-01-01

The main emphasis is upon the propagation angle with respect to the duct axis and its relation to the far-field acoustic radiation pattern. When the steady flow Mach number is accounted for in the duct, the propagation angle in the duct is shown to be coincident with the angle of the principal lobe of far-field radiation obtained using the Wiener-Hopf technique. Different Mach numbers are allowed within the duct and in the external field. For static tests with a steady flow in an inlet but with no external Mach number the far-field radiation pattern is shifted considerably toward the inlet axis when compared to zero Mach number radiation theory. As the external Mach number is increased the noise radiation pattern is shifted away from the inlet axis. The theory is developed using approximations for sound propagation in circular ducts. An exact analysis using Hankel function solutions for the zero Mach number case is given to provide a check of the simpler approximate theory.

Wicket gate trailing-edge blowing: A method for improving off-design hydroturbine performance by adjusting the runner inlet swirl angle

NASA Astrophysics Data System (ADS)

Lewis, B. J.; Cimbala, J. M.; Wouden, A. M.

2014-03-01

At their best efficiency point (BEP), hydroturbines operate at very high efficiency. However, with the ever-increasing penetration of alternative electricity generation, it has become common to operate hydroturbines at off-design conditions in order to maintain stability in the electric power grid. This paper demonstrates a method for improving hydroturbine performance during off-design operation by injecting water through slots at the trailing edges of the wicket gates. The injected water causes a change in bulk flow direction at the inlet of the runner. This change in flow angle from the wicket gate trailing-edge jets provides the capability of independently varying the flow rate and swirl angle through the runner, which in current designs are both determined by the wicket gate opening angle. When properly tuned, altering the flow angle results in a significant improvement in turbine efficiency during off-design operation.

Evaluation of F/A-18A HARV inlet flow analysis with flight data

NASA Technical Reports Server (NTRS)

Smith, C. Frederic; Podleski, Steve D.; Barankiewicz, Wendy S.; Zeleznik, Susan Z.

1995-01-01

The F/A-18A aircraft has experienced engine stalls at high angles-of-attack and yaw flight conditions which were outside of its flight envelope. Future aircraft may be designed to operate routinely in this flight regime. Therefore, it is essential that an understanding of the inlet flow field at these flight conditions be obtained. Due to the complex interactions of the fuselage and inlet flow fields, a study of the flow within the inlet must also include external effects. Full Navier-Stokes (FNS) calculations on the F/A-18A High Alpha Research Vehicle (HARV) inlet for several angles-of-attack with sideslip and free stream Mach numbers have been obtained. The predicted forebody/fuselage surface static pressures agreed well with flight data. The surface static pressures along the inlet lip are in good agreement with the numerical predictions. The major departure in agreement is along the bottom of the lip at 30 deg and 60 deg angle-of-attack where a possible streamwise flow separation is not being predicted by the code. The circumferential pressure distributions at the engine face are in very good agreement with the numerical results. The variation in surface static pressure in the circumferential direction is very small with the exception of 60 angle-of-attack. Although the simulation does not include the effect of the engine, it appears that this omission has a second order effect on the circumferential pressure distribution. An examination of the unsteady flight test data base has shown that the secondary vortex migrates a significant distance with time. In fact, the extent of this migration increases with angle-of-attack with increasing levels of distortion. The effects of the engine on this vortex movement is unknown. This implies that the level of flow unsteadiness increases with increasing distortion. Since the computational results represent an asymptotic solution driven by steady boundary conditions, these numerical results may represent an arbitrary point in time. A comparison of the predicted total pressure contours with flight data indicates that the numerical results are within the excursion range of the unsteady data which is the best the calculations can attain unless an unsteady simulation is performed.

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.

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.

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.

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.

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.

Optimal operation of turbo blowers serially connected using inlet vanes

NASA Astrophysics Data System (ADS)

Jang, Choon-Man

2011-03-01

Optimal operation of the turbo blowers having an inlet vane has been studied to understand the blowers' operating performance. To analyze three-dimensional flow field in the turbo blowers serially connected, general analysis code, CFX, is introduced in the present work. SST turbulence model is employed to estimate the eddy viscosity. Throughout the numerical analysis, it is found that the flow rates of the turbo blowers can be controlled at the vane angle between 90 (full open condition) degrees and 60 degrees effectively, because pressure loss rapidly increases below 60 degree of a vane angle. Efficiency also has almost the same values from 90 degrees to 60 degrees of a vane angle. It is noted that the distorted inlet velocity generated in the small vane angle makes performance deterioration of the turbo blowers due to the local leading edge separation and the following non-uniform blade loading.

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.

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

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.

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.

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

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.

Some Effects of Compressibility on the Flow Through Fans and Turbines

NASA Technical Reports Server (NTRS)

Perl, W.; Epstein, H. T.

1946-01-01

The laws of conservation of mass, momentum, and energy are applied to the compressible flow through a two-dimensional cascade of airfoils. A fundamental relation between the ultimate upstream and downstream flow angles, the inlet Mach number, and the pressure ratio across the cascade is derived. Comparison with the corresponding relation for incompressible flow shows large differences. The fundamental relation reveals two ranges of flow angles and inlet Mach numbers, for which no ideal pressure ratio exists. One of these nonideal operating ranges is analogous to a similar type in incompressible flow. The other is characteristic only of compressible flow. The effect of variable axial-flow area is treated. Some implications of the basic conservation laws in the case of nonideal flow through cascades are discussed.

Low speed tests of a fixed geometry inlet for a tilt nacelle V/STOL airplane

NASA Technical Reports Server (NTRS)

Syberg, J.; Koncsek, J. L.

1977-01-01

Test data were obtained with a 1/4 scale cold flow model of the inlet at freestream velocities from 0 to 77 m/s (150 knots) and angles of attack from 45 deg to 120 deg. A large scale model was tested with a high bypass ratio turbofan in the NASA/ARC wind tunnel. A fixed geometry inlet is a viable concept for a tilt nacelle V/STOL application. Comparison of data obtained with the two models indicates that flow separation at high angles of attack and low airflow rates is strongly sensitive to Reynolds number and that the large scale model has a significantly improved range of separation-free operation.

Effect of inlet cone pipe angle in catalytic converter

NASA Astrophysics Data System (ADS)

Amira Zainal, Nurul; Farhain Azmi, Ezzatul; Arifin Samad, Mohd

2018-03-01

The catalytic converter shows significant consequence to improve the performance of the vehicle start from it launched into production. Nowadays, the geometric design of the catalytic converter has become critical to avoid the behavior of backpressure in the exhaust system. The backpressure essentially reduced the performance of vehicles and increased the fuel consumption gradually. Consequently, this study aims to design various models of catalytic converter and optimize the volume of fluid flow inside the catalytic converter by changing the inlet cone pipe angles. Three different geometry angles of the inlet cone pipe of the catalytic converter were assessed. The model is simulated in Solidworks software to determine the optimum geometric design of the catalytic converter. The result showed that by decreasing the divergence angle of inlet cone pipe will upsurge the performance of the catalytic converter.

Force and Pressure Recovery Characteristics at Supersonic Speeds of a Conical Spike Inlet with a Bypass Discharging from the Top or Bottom of the Diffuser in an Axial Direction

NASA Technical Reports Server (NTRS)

Allen, J L; Beke, Andrew

1953-01-01

Force and pressure-recovery characteristics of a nacelle-type conical-spike inlet with a fixed-area bypass located in the top or bottom of the diffuser are presented for flight Mach numbers of 1.6, 1.8, and 2.0 for angles of attack from 0 degrees to 9 degrees. Top or bottom location of the bypass did not have significant effects on diffuser pressure-recovery, bypass mass-flow ratio, or drag coefficient over the range of angles of attack, flight Mach numbers, and stable engine mass-flow ratios investigated. A larger stable subcritical operating range was obtained with the bypass on the bottom at angles of attack from 3 degrees to 9 degrees at a flight Mach number of 2.0. At a flight Mach number of 2.0, the discharge of 14 percent of the critical mass flow of the inlet by means of a bypass increased the drag only one-fifth of the additive drag that would result for equivalent spillage behind an inlet normal shock without significant reductions in diffuser pressure recovery.

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.

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.

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.

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.

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.

Automatic efficiency optimization of an axial compressor with adjustable inlet guide vanes

NASA Astrophysics Data System (ADS)

Li, Jichao; Lin, Feng; Nie, Chaoqun; Chen, Jingyi

2012-04-01

The inlet attack angle of rotor blade reasonably can be adjusted with the change of the stagger angle of inlet guide vane (IGV); so the efficiency of each condition will be affected. For the purpose to improve the efficiency, the DSP (Digital Signal Processor) controller is designed to adjust the stagger angle of IGV automatically in order to optimize the efficiency at any operating condition. The A/D signal collection includes inlet static pressure, outlet static pressure, outlet total pressure, rotor speed and torque signal, the efficiency can be calculated in the DSP, and the angle signal for the stepping motor which control the IGV will be sent out from the D/A. Experimental investigations are performed in a three-stage, low-speed axial compressor with variable inlet guide vanes. It is demonstrated that the DSP designed can well adjust the stagger angle of IGV online, the efficiency under different conditions can be optimized. This establishment of DSP online adjustment scheme may provide a practical solution for improving performance of multi-stage axial flow compressor when its operating condition is varied.

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.

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.

PAN AIR application to the F-106B

NASA Technical Reports Server (NTRS)

Ghaffari, F.

1986-01-01

The PAN AIR computer code was employed in the present study to investigate the aerodynamic effects of the various geometrical changes and flow conditions on a configuration similar to the F-106B half-airplane tested in the Langley 30x60-foot wind tunnel. The various geometries studied included two forebodies (original and shortened), two inlet flow conditions (open and closed) two vortex flap situations (off and on). The attached flow theoretical solutions were obtained for Mach number of 0.08 and angle of attack of 8 deg., 10 deg., 12 deg., and 14 deg. In general this investigation revealed that the shortening of the forebody or closing of the inlet produced only a small change in the overall aerodynamic coefficients of the basic F-106B configuration throughout the examined angles of attack. However, closing the inlet of the configuration resulted in a slightly higher drag level at low angles of attack. Furthermore, at and above 10 deg. angle of attack, it was shown that the presence of the vortex flap causes an increase in the total lift and drag. Also, these theoretical results showed the expected reduction in longitudinal stability level with addition of the vortex flap to the basic F-106B configuration.

Flight and wind-tunnel comparisons of the inlet-airframe interaction of the F-15 airplane

NASA Technical Reports Server (NTRS)

Webb, L. D.; Andriyich-Varda, D.; Whitmore, S. A.

1984-01-01

The design of inlets and nozzles and their interactions with the airplane which may account for a large percentage of the total drag of modern high performance aircraft is discussed. The inlet/airframe interactions program and the flight tests conducted is described. Inlet drag and lift data from a 7.5% wind-tunnel model are compared with data from an F-15 airplane with instrumentation to match the model. Pressure coefficient variations with variable cowl angles, capture ratios, examples of flow interactions and angles of attack are for Mach numbers of 0.6, 0.9, 1.2, and 1.5 are presented.

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2013-01-01

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×10(exp 5) to 0.85×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade. Revision 1

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2014-01-01

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50 deg or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×10 (exp 5) to 0.85×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2012-01-01

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented. Measurements were made in NASA's Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50% speed range from takeoff to altitude cruise. This results in 50 degrees or more variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83 × 10(exp 5) to 0.85 ×10(exp 5) and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane 8.6% axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition

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.

Cold-air performance of compressor-drive turbine of Department of Energy upgraded automobile gas turbine engine. 1: Volute-manifold and stator performance

NASA Technical Reports Server (NTRS)

Roelke, R. J.; Haas, J. E.

1981-01-01

The aerodynamic performance of the inlet manifold and stator assembly of the compressor drive turbine was experimentally determined with cold air as the working fluid. The investigation included measurements of mass flow and stator-exit fluid torque as well as radial surveys of total pressure and flow angle at the stator inlet and annulus surveys of total pressure and flow angle at the stator exit. The stator-exit aftermixed flow conditions and overall stator efficiency were obtained and compared with their design values and the experimental results from three other stators. In addition, an analysis was made to determine the constituent aerodynamic losses that made up the stator kinetic energy loss.

Powder collection apparatus/method

DOEpatents

Anderson, I.E.; Terpstra, R.L.; Moore, J.A.

1994-01-11

Device for separating and collecting ultrafine atomized powder from the gas stream of a gas atomizing apparatus comprises a housing having an interior wall oriented at an angle relative to horizontal so as to form a downwardly converging, conical expansion chamber, an inlet conduit communicated to the expansion chamber proximate an upper region thereof for receiving the gas stream, and an outlet proximate a lower region of the expansion chamber. The inlet conduit is oriented at a compound inclined angle (with respect to horizontal) selected to promote separation and collection of powder from the gas stream in the expansion chamber. The compound angle comprises a first entrance angle that is greater than the angle of repose of the powder on the housing interior wall such that any powder accumulation in the inlet conduit tends to flow down the wall toward the outlet. The second angle is selected generally equal to the angle of the housing interior wall measured from the same horizontal plane so as to direct the gas stream into the expansion chamber generally tangent to the housing interior wall to establish a downward swirling gas stream flow in the expansion chamber. A powder collection container is communicated to the outlet of the expansion chamber to collect the powder for further processing. 4 figures.

Powder collection apparatus/method

DOEpatents

Anderson, Iver E.; Terpstra, Robert L.; Moore, Jeffery A.

1994-01-11

Device for separating and collecting ultrafine atomized powder from the gas stream of a gas atomizing apparatus comprises a housing having an interior wall oriented at an angle relative to horizontal so as to form a downwardly converging, conical expansion chamber, an inlet conduit communicated to the expansion chamber proximate an upper region thereof for receiving the gas stream, and an outlet proximate a lower region of the expansion chamber. The inlet conduit is oriented at a compound inclined angle (with respect to horizontal) selected to promote separation and collection of powder from the gas stream in the expansion chamber. The compound angle comprises a first entrance angle that is greater than the angle of repose of the powder on the housing interior wall such that any powder accumulation in the inlet conduit tends to flow down the wall toward the outlet. The second angle is selected generally equal to the angle of the housing interior wall measured from the same horizontal plane so as to direct the gas stream into the expansion chamber generally tangent to the housing interior wall to establish a downward swirling gas stream flow in the expansion chamber. A powder collection container is communicated to the outlet of the expansion chamber to collect the powder for further processing.

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.

Supersonic flow visualization of a nacelle in close proximity to a simulated wing

NASA Technical Reports Server (NTRS)

Biber, Kasim; Ellis, David R.

1993-01-01

A flow visualization study was made in the 9 x 9 inch supersonic wind tunnel at Wichita State University to examine shock and boundary layer flow interaction for a nacelle in close proximity to the lower surface of a simulated wing. The test matrix included variations of angle of attack from -2 degrees to +4 degrees, nacelle-wing gap from 0.5 to 3-nacelle inlet diameter (0.12 inch), and Reynolds number based on nacelle length (1.164 inch) from 1.16 x 10(exp 6) to 1.45 x 10(exp 6) at a nominal Mach number of 2. Schlieren pictures of wing and nacelle flowfield were recorded by a video camera during each tunnel run. Results show that the nacelle inlet shock wave remains attached to the inlet lip and its impingement does not significantly affect the wing boundary layer. At the nacelle trailing edge location, the wing boundary layer thickness is approximately one nacelle inlet diameter at alpha = 0 degrees and it decreases with increase of angle of attack.

Inlet Flow Characteristics During Rapid Maneuvers for an F/A-18A Airplane

NASA Technical Reports Server (NTRS)

Steenken, William G.; Williams, John G.; Walsh, Kevin R.

1999-01-01

The F404-GE-400 engine powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the characteristics of inlet airflow during rapid aircraft maneuvers. A study of the degree of similarity between inlet data obtained during rapid aircraft maneuvers and inlet data obtained at steady aerodynamic attitudes was conducted at the maximum engine airflow of approximately 145 Ibm/sec using a computer model that was generated from inlet data obtained during steady aerodynamic maneuvers. Results show that rapid-maneuver inlet recoveries agreed very well with the recoveries obtained at equivalent stabilized angle-of-attack conditions. The peak dynamic circumferential distortion values obtained during rapid maneuvers agreed within 0.01 units of distortion over the 10 - 38 degree angle of attack range with the values obtained during steady aerodynamic maneuvers while similar agreement was found for the peak dynamic radial distortion values up to 29 degrees angle-of-attack. Exceedences of the rapid-maneuver peak dynamic circumferential distortion values relative to the peak distortion model values at steady attitudes occurred only at low or negative angles of attack and were inconsequential from an engine-stability assessment point of view. The results of this study validate the current industry practice of testing at steady aerodynamic conditions to characterize inlet recovery and peak dynamic distortion levels.

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.

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

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.

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.

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.

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.

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.

Acoustic Signatures of a Model Fan in the NASA-Lewis Anechoic Wind Tunnel

NASA Technical Reports Server (NTRS)

Dietrich, D. A.; Heidmann, M. F.; Abbott, J. M.

1977-01-01

One-third octave band and narrowband spectra and continuous directivity patterns radiated from an inlet are presented over ranges of fan operating conditions, tunnel velocity, and angle of attack. Tunnel flow markedly reduced the unsteadiness and level of the blade passage tone, revealed the cutoff design feature of the blade passage tone, and exposed a lobular directivity pattern for the second harmonic tone. The full effects of tunnel flow are shown to be complete above a tunnel velocity of 20 meters/second. The acoustic signatures are also shown to be strongly affected by fan rotational speed, fan blade loading, and inlet angle of attack.

Numerical simulation of supersonic and hypersonic inlet flow fields

NASA Technical Reports Server (NTRS)

Mcrae, D. Scott; Kontinos, Dean A.

1995-01-01

This report summarizes the research performed by North Carolina State University and NASA Ames Research Center under Cooperative Agreement NCA2-719, 'Numerical Simulation of Supersonic and Hypersonic Inlet Flow Fields". Four distinct rotated upwind schemes were developed and investigated to determine accuracy and practicality. The scheme found to have the best combination of attributes, including reduction to grid alignment with no rotation, was the cell centered non-orthogonal (CCNO) scheme. In 2D, the CCNO scheme improved rotation when flux interpolation was extended to second order. In 3D, improvements were less dramatic in all cases, with second order flux interpolation showing the least improvement over grid aligned upwinding. The reduction in improvement is attributed to uncertainty in determining optimum rotation angle and difficulty in performing accurate and efficient interpolation of the angle in 3D. The CCNO rotational technique will prove very useful for increasing accuracy when second order interpolation is not appropriate and will materially improve inlet flow solutions.

Standardized solar simulator tests of flat plate solar collectors. 1: Soltex collector with two transparent covers

NASA Technical Reports Server (NTRS)

Simon, F.

1975-01-01

A Soltex flat plate solar collector was tested with a solar simulator for inlet temperatures of 77 to 201 F, flux levels of 240 and 350 Btu/hr-sq ft, a collant flow rate of 10.5 lb/hr sq ft, and incident angles of 0 deg, 41.5 deg, and 65.2 deg. Collector performance is correlated in terms of inlet temperature, flux level, and incident angle.

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.

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.

Low-speed wind tunnel tests of a 50.8-centimeter (20-in.) 1.15-pressure-ratio fan engine model

NASA Technical Reports Server (NTRS)

Wesoky, H. L.; Abbott, J. M.; Albers, J. A.; Dietrich, D. A.

1974-01-01

At a typical STOL aircraft takeoff and landing velocity, wind tunnel aerodynamic and acoustic measurements demonstrated that an inlet lip-area contraction ratio of 1.35 was superior to a ratio of 1.26 at high incidence angles. A 17 percent reduction in net thrust and an increase of 9 decibels in sound pressure level at the blade passing frequency resulted from inlet flow separation at an incidence angle of 50 deg with the 1.26-contraction-ratio inlet. Reverse-thrust forces obtained with blade rotation through the feathered angle were 1.8 times larger than with blade rotation through the flat angle. Reverse-thrust force was reduced from 30 to 50 percent and sound pressure level increased from 3 to 7 decibels at the blade passing frequency between the wind-tunnel-off condition and a typical STOL aircraft landing velocity.

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.

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.

Flight-determined characteristics of an air intake system on an F-111A airplane

NASA Technical Reports Server (NTRS)

Hughes, D. L.; Johnson, H. J.

1972-01-01

Flow phenomena of the F-111A air intake system were investigated over a large range of Mach number, altitude, and angle of attack. Boundary-layer variations are shown for the fuselage splitter plate and inlet entrance stations. Inlet performance is shown in terms of pressure recovery, airflow, mass-flow ratio, turbulence factor, distortion factor, and power spectral density. The fuselage boundary layer was found to be not completely removed from the upper portion of the splitter plate at all Mach numbers investigated. Inlet boundary-layer ingestion started at approximately Mach 1.6 near the translating spike and cone. Pressure-recovery distribution at the compressor face showed increasing distortion with increasing angle of attack and increasing Mach number. The time-averaged distortion-factor value approached 1300, which is near the distortion tolerance of the engine at Mach numbers above 2.1.

Calculation of the flow field including boundary layer effects for supersonic mixed compression inlets at angles of attack

NASA Technical Reports Server (NTRS)

Vadyak, J.; Hoffman, J. D.

1982-01-01

The flow field in supersonic mixed compression aircraft inlets at angle of attack is calculated. A zonal modeling technique is employed to obtain the solution which divides the flow field into different computational regions. The computational regions consist of a supersonic core flow, boundary layer flows adjacent to both the forebody/centerbody and cowl contours, and flow in the shock wave boundary layer interaction regions. The zonal modeling analysis is described and some computational results are presented. The governing equations for the supersonic core flow form a hyperbolic system of partial differential equations. The equations for the characteristic surfaces and the compatibility equations applicable along these surfaces are derived. The characteristic surfaces are the stream surfaces, which are surfaces composed of streamlines, and the wave surfaces, which are surfaces tangent to a Mach conoid. The compatibility equations are expressed as directional derivatives along streamlines and bicharacteristics, which are the lines of tangency between a wave surface and a Mach conoid.

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.

Iso-thermal flow characteristics of rotationally symmetric jets generating a swirl within a cylindrical chamber

NASA Astrophysics Data System (ADS)

Long, Shen; Lau, Timothy C. W.; Chinnici, Alfonso; Tian, Zhao Feng; Dally, Bassam B.; Nathan, Graham J.

2018-05-01

We present a systematic experimental study of the interaction between four rotationally symmetric jets within a cylindrical chamber, under conditions relevant to a wide range of engineering applications, including the technology of a Hybrid Solar Receiver Combustor (HSRC). The HSRC geometry is simplified here to a cylindrical cavity with four inlet jets (representing four burners) which are configured in an annular arrangement and aligned at an inclination angle to the axis with a tangential component (azimuthal angle) to generate a swirl in the chamber. In this study, the jet inclination angle (αj) was varied over the range of 25°-45°, while the jet azimuthal angle (θj) was varied from 5° to 15°. The inlet Reynolds number for each injected jet and the number of jets were fixed at ReD = 10 500 and 4, respectively. Measurements obtained with Particle Image Velocimetry were used to characterise the large-scale flow field within selected configurations. The results reveal a significant dependence of the mean and root-mean-square flow-fields on the jet azimuthal angle (θj) and the jet inclination angle (αj). Three different flow regimes with distinctive flow characteristics were identified within the configurations investigated here. It was also found that θj can significantly influence (a) the position and strength of an external recirculation zone and a central recirculation zone, (b) the extent of turbulence fluctuation, and (c) the flow unsteadiness. Importantly, the effect of αj on the flow characteristics was found to depend strongly on the value of θj.

A computer program for the calculation of the flow field including boundary layer effects for mixed-compression inlets at angle of attack

NASA Technical Reports Server (NTRS)

Vadyak, J.; Hoffman, J. D.

1982-01-01

A computer program was developed which is capable of calculating the flow field in the supersonic portion of a mixed compression aircraft inlet operating at angle of attack. The supersonic core flow is computed using a second-order three dimensional method-of-characteristics algorithm. The bow shock and the internal shock train are treated discretely using a three dimensional shock fitting procedure. The boundary layer flows are computed using a second-order implicit finite difference method. The shock wave-boundary layer interaction is computed using an integral formulation. The general structure of the computer program is discussed, and a brief description of each subroutine is given. All program input parameters are defined, and a brief discussion on interpretation of the output is provided. A number of sample cases, complete with data listings, are provided.

Predictions and measurements of isothermal flowfields in axisymmetric combustor geometries. Ph.D. Thesis. Final Report

NASA Technical Reports Server (NTRS)

Rhodes, D. L.; Lilley, D. G.

1985-01-01

Numerical predictions, flow visualization experiments and time-mean velocity measurements were obtained for six basic nonreacting flowfields (with inlet swirl vane angles of 0 (swirler removed), 45 and 70 degrees and sidewall expansion angles of 90 and 45 degrees) in an idealized axisymmetric combustor geometry. A flowfield prediction computer program was developed which solves appropriate finite difference equations including a conventional two equation k-epsilon eddy viscosity turbulence model. The wall functions employed were derived from previous swirling flow measurements, and the stairstep approximation was employed to represent the sloping wall at the inlet to the test chamber. Recirculation region boundaries have been sketched from the entire flow visualization photograph collection. Tufts, smoke, and neutrally buoyant helium filled soap bubbles were employed as flow tracers. A five hole pitot probe was utilized to measure the axial, radial, and swirl time mean velocity components.

Unified Model for the Overall Efficiency of Inlets Sampling from Horizontal Aerosol Flows

NASA Astrophysics Data System (ADS)

Hangal, Sunil Pralhad

When sampling aerosols from ambient or industrial air environments, the sampled aerosol must be representative of the aerosol in the free stream. The changes that occur during sampling must be assessed quantitatively so that sampling errors can be compensated for. In this study, unified models have been developed for the overall efficiency of tubular sharp-edged inlets sampling from horizontal aerosol flows oriented at 0 to 90^circ relative to the wind direction in the vertical (pitch) and horizontal plane(yaw). In the unified model, based on experimental data, the aspiration efficiency is represented by a single equation with different inertial parameters at 0 to 60^ circ and 45 to 90^circ . Tnt transmission efficiency is separated into two components: one due to gravitational settling in the boundary layer and the other due to impaction. The gravitational settling component is determined by extending a previously developed isoaxial sampling model to nonisoaxial sampling. The impaction component is determined by a new model that quantifies the particle losses caused by wall impaction. The model also quantifies the additional particle losses resulting from turbulent motion in the vena contracta which is formed in the inlet when the inlet velocity is higher than the wind velocity. When sampling aerosols in ambient or industrial environments with an inlet, small changes in wind direction or physical constraints in positioning the inlet in the system necessitates the assessment of sampling efficiency in both the vertical and horizontal plane. The overall sampling efficiency of tubular inlets has been experimentally investigated in yaw and pitch orientations at 0 to 20 ^circ from horizontal aerosol flows using a wind tunnel facility. The model for overall sampling efficiency has been extended to include both yaw and pitch sampling based on the new data. In this model, the difference between yaw and pitch is expressed by the effect of gravity on the impaction process inside the inlet described by a newly developed gravity effect angle. At yaw, the gravity effect angle on the wall impaction process does not change with sampling angle. At pitch, the gravity effect on the impaction process results in particle loss increase for upward and decrease for downward sampling. Using the unified model, graphical representations have been developed for sampling at small angles. These can be used in the field to determine the overall sampling efficiency of inlets at several operating conditions and the operating conditions that result in an acceptable sampling error. Pitch and diameter factors have been introduced for relating the efficiency values over a wide range of conditions to those of a reference condition. The pitch factor determines the overall sampling efficiency at pitch from yaw values, and the diameter factor determines the overall sampling efficiency at different inlet diameters.

Two-Dimensional Supersonic Nozzle Thrust Vectoring Using Staggered Ramps

NASA Astrophysics Data System (ADS)

Montes, Carlos Fernando

A novel mechanism for vectoring the thrust of a supersonic, air-breathing engine was analyzed numerically using ANSYS Fluent. The mechanism uses two asymmetrically staggered ramps; one placed at the throat, the other positioned at the exit lip of the nozzle. The nozzle was designed using published flow data, isentropic relationships, and piecewise quartic splines. The design was verified numerically and was in fair agreement with the analytical data. Using the steady-state pressure-based solver, along with the realizable kappa - epsilon turbulence model, a total of eighteen simulations were conducted: three ramp lengths at three angles, using two sets of inlet boundary conditions (non-afterburning and afterburning). The vectoring simulations showed that the afterburning flow yields a lower flow deflection distribution, shown by the calculated average deflection angle and area-weighted integrals of the distributions. The data implies that an aircraft can achieve an average thrust vectoring angle of approximately 30° in a given direction with the longest ramp length and largest ramp angle configuration. With increasing ramp angle, the static pressure across the nozzle inlet increased, causing concern for potential negative effects on the engine's turbine. The mechanism, for which a provisional patent application has been filed, will require further work to investigate the maximum possible thrust vectoring angle, including experiments.

Inlet Turbulence and Length Scale Measurements in a Large Scale Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Thurman, Douglas; Flegel, Ashlie; Giel, Paul

2014-01-01

Constant temperature hotwire anemometry data were acquired to determine the inlet turbulence conditions of a transonic turbine blade linear cascade. Flow conditions and angles were investigated that corresponded to the take-off and cruise conditions of the Variable Speed Power Turbine (VSPT) project and to an Energy Efficient Engine (EEE) scaled rotor blade tip section. Mean and turbulent flowfield measurements including intensity, length scale, turbulence decay, and power spectra were determined for high and low turbulence intensity flows at various Reynolds numbers and spanwise locations. The experimental data will be useful for establishing the inlet boundary conditions needed to validate turbulence models in CFD codes.

Transition aerodynamics for 20-percent-scale VTOL unmanned aerial vehicle

NASA Technical Reports Server (NTRS)

Kjerstad, Kevin J.; Paulson, John W., Jr.

1993-01-01

An investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to establish a transition data base for an unmanned aerial vehicle utilizing a powered-lift ejector system and to evaluate alterations to the ejector system for improved vehicle performance. The model used in this investigation was a 20-percent-scale, blended-body, arrow-wing configuration with integrated twin rectangular ejectors. The test was conducted from hover through transition conditions with variations in angle of attack, angle of sideslip, free-stream dynamic pressure, nozzle pressure ratio, and model ground height. Force and moment data along with extensive surface pressure data were obtained. A laser velocimeter technique for measuring inlet flow velocities was demonstrated at a single flow condition, and also a low order panel method was successfully used to numerically simulate the ejector inlet flow.

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.

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.

A Robust Design Methodology for Optimal Microscale Secondary Flow Control in Compact Inlet Diffusers

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Keller, Dennis J.

2001-01-01

It is the purpose of this study to develop an economical Robust design methodology for microscale secondary flow control in compact inlet diffusers. To illustrate the potential of economical Robust Design methodology, two different mission strategies were considered for the subject inlet, namely Maximum Performance and Maximum HCF Life Expectancy. The Maximum Performance mission maximized total pressure recovery while the Maximum HCF Life Expectancy mission minimized the mean of the first five Fourier harmonic amplitudes, i.e., 'collectively' reduced all the harmonic 1/2 amplitudes of engine face distortion. Each of the mission strategies was subject to a low engine face distortion constraint, i.e., DC60<0.10, which is a level acceptable for commercial engines. For each of these missions strategies, an 'Optimal Robust' (open loop control) and an 'Optimal Adaptive' (closed loop control) installation was designed over a twenty degree angle-of-incidence range. The Optimal Robust installation used economical Robust Design methodology to arrive at a single design which operated over the entire angle-of-incident range (open loop control). The Optimal Adaptive installation optimized all the design parameters at each angle-of-incidence. Thus, the Optimal Adaptive installation would require a closed loop control system to sense a proper signal for each effector and modify that effector device, whether mechanical or fluidic, for optimal inlet performance. In general, the performance differences between the Optimal Adaptive and Optimal Robust installation designs were found to be marginal. This suggests, however, that Optimal Robust open loop installation designs can be very competitive with Optimal Adaptive close loop designs. Secondary flow control in inlets is inherently robust, provided it is optimally designed. Therefore, the new methodology presented in this paper, combined array 'Lower Order' approach to Robust DOE, offers the aerodynamicist a very viable and economical way of exploring the concept of Robust inlet design, where the mission variables are brought directly into the inlet design process and insensitivity or robustness to the mission variables becomes a design objective.

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.

Performance of a Supersonic Over-Wing Inlet with Application to a Low-Sonic-Boom Aircraft

NASA Technical Reports Server (NTRS)

Trefny, Charles J.; Hirt, Stefanie M.; Anderson, Bernhard H.; Fink, Lawrence E.; Magee, Todd E.

2014-01-01

Development of commercial supersonic aircraft has been hindered by many related factors including fuel-efficiency, economics, and sonic-boom signatures that have prevented over-land flight. Materials, propulsion, and flight control technologies have developed to the point where, if over-land flight were made possible, a commercial supersonic transport could be economically viable. Computational fluid dynamics, and modern optimization techniques enable designers to reduce the boom signature of candidate aircraft configurations to acceptable levels. However, propulsion systems must be carefully integrated with these low-boom configurations in order that the signatures remain acceptable. One technique to minimize the downward propagation of waves is to mount the propulsion systems above the wing, such that the wing provides shielding from shock waves generated by the inlet and nacelle. This topmounted approach introduces a number of issues with inlet design and performance especially with the highly-swept wing configurations common to low-boom designs. A 1.79%-scale aircraft model was built and tested at the NASA Glenn Research Center's 8-by 6-Foot Supersonic Wind Tunnel (8x6 SWT) to validate the configuration's sonic boom signature. In order to evaluate performance of the top-mounted inlets, the starboard flow-through nacelle on the aerodynamic model was replaced by a 2.3%-scale operational inlet model. This integrated configuration was tested at the 8x6 SWT from Mach 0.25 to 1.8 over a wide range of angles-of-attack and yaw. The inlet was also tested in an isolated configuration over a smaller range of angles-of-attack and yaw. A number of boundary-layer bleed configurations were investigated and found to provide a substantial positive impact on pressure recovery and distortion. Installed inlet performance in terms of mass capture, pressure recovery, and distortion over the Mach number range at the design angle-of-attack of 4-degrees is presented herein and compared to that at 0- degrees, as well as the isolated inlet configuration to highlight installation effects. Performance of the installed inlet fell below that of the isolated inlet at Mach numbers of 1.4 and greater. The installed inlet demonstrated adequate operability over the expected range of angles-of-attack and yaw, but did exhibit definite angle-ofattack and yaw limits at supersonic conditions. At each supersonic flight Mach number, performance parameters near zero yaw angle were relatively insensitive to yaw, but in general the yaw angle yielding best performance was non-zero and varied with angle-of-attack. Performance of the installed inlet is also presented as functions of angle-of-attack and yaw to highlight these effects. Distortion at the aerodynamic interface plane ranged between 10 and 25% at the inlet critical points over the range of flight Mach numbers tested and did not decrease significantly for the isolated inlet. Although these distortion levels would be considered high for operation with a turbine engine, the over-wing installation is likely not as significant a contributor as the low test Reynolds number. This is demonstrated by comparing CFD analysis of the isolated inlet at test scale with that at intermediate and full scales.

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.

Throat-bypass bleed systems for increasing the stable airflow range of a Mach 2.50 axisymmetric inlet with 40-percent internal contraction

NASA Technical Reports Server (NTRS)

Sanders, B. W.; Mitchell, G. A.

1973-01-01

The results of an experimental investigation to increase the stable airflow range of a super sonic mixed-compression inlet are presented. Various throat-bypass bleeds were located on the inlet cowl. The bleed types were distributed porous normal holes, a forward slanted slot, or distributed educated slots. Large inlet stability margins were obtained with the inlet throat bleed systems if a constant pressure was maintained in the throat-bypass bleed plenum. Stability limits were determined for steady-state and limited transient internal air flow changes. Limited unstart angle-of-attack data are presented.

Comparison of experimental and theoretical boundary-layer separation for inlets at incidence angle at low-speed conditions

NASA Technical Reports Server (NTRS)

Felderman, E. J.; Albers, J. A.

1975-01-01

Comparisons between experimental and theoretical Mach number distributions and separation locations are presented for the internal surfaces of four different subsonic inlet geometries with exit diameters of 13.97 centimeters. The free stream Mach number was held constant at 0.127, the one-dimensional throat Mach number ranged from 0.49 to 0.71, and the incidence angle ranged from 0 deg to 50 deg. Generally good agreement was found between the theoretical and experimental surface Mach number distributions as long as no flow separation existed. At high incidence angles, where separation was obvious in the experimental data, the theory predicted separation on the lip. At lower incidence angles, the theoretical results indicated diffuser separation which was not obvious from the experimental surface Mach number distributions. As incidence angle was varied from 0 deg to 50 deg, the predicted separation location shifted from the diffuser region to the inlet highlight. Relatively small total pressure losses were obtained when the predicted separation location was greater than 0.6 of the distance between the highlight and the diffuser exit.

Low-speed aerodynamic test of an axisymmetric supersonic inlet with variable cowl slot

NASA Technical Reports Server (NTRS)

Powell, A. G.; Welge, H. R.; Trefny, C. J.

1985-01-01

The experimental low-speed aerodynamic characteristics of an axisymmetric mixed-compression supersonic inlet with variable cowl slot are described. The model consisted of the NASA P-inlet centerbody and redesigned cowl with variable cowl slot powered by the JT8D single-stage fan simulator and driven by an air turbine. The model was tested in the NASA Lewis Research Center 9- by 15-foot low-speed tunnel at Mach numbers of 0, 0.1, and 0.2 over a range of flows, cowl slot openings, centerbody positions, and angles of attack. The variable cowl slot was effective in minimizing lip separation at high velocity ratios, showed good steady-state and dynamic distortion characteristics, and had good angle-of-attack tolerance.

New centrifugal blood pump with dual impeller and double pivot bearing system: wear evaluation in bearing system, performance tests, and preliminary hemolysis tests.

PubMed

Bock, Eduardo; Ribeiro, Adriana; Silva, Maxwell; Antunes, Pedro; Fonseca, Jeison; Legendre, Daniel; Leme, Juliana; Arruda, Celso; Biscegli, José; Nicolosi, Denys; Andrade, Aron

2008-04-01

A new dual impeller centrifugal blood pump has been developed as a research collaboration between Baylor College of Medicine and Institute Dante Pazzanese of Cardiology for long-term left ventricle assist device (LVAD). A design feature of this new pump is a dual impeller that aims to minimize a stagnant flow pattern around the inlet port. Several different materials were tested in order to adopt a double pivot bearing design originally developed by Prof. Dr. Yukihiko Nosé from Baylor College of Medicine. Hydraulic performance tests were conducted with two different inlet ports' angle configurations 30 degrees and 45 degrees . Pump with inlet port angle of 45 degrees achieved best values of pressure ahead and flow after 1800 rpm. Preliminary hemolysis tests were conducted using human blood. The pump showed good performance results and no alarming trace of hemolysis, proving to be a feasible long-term LVAD.

Computational Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Takeoff and Landing Conditions

NASA Technical Reports Server (NTRS)

Tompkins, Daniel M.; Sexton, Matthew R.; Mugica, Edward A.; Beyar, Michael D.; Schuh, Michael J.; Stremel, Paul M.; Deere, Karen A.; McMillin, Naomi; Carter, Melissa B.

2016-01-01

Due to the aft, upper surface engine location on the Hybrid Wing Body (HWB) planform, there is potential to shed vorticity and separated wakes into the engine when the vehicle is operated at off-design conditions and corners of the envelope required for engine and airplane certification. CFD simulations were performed of the full-scale reference propulsion system, operating at a range of inlet flow rates, flight speeds, altitudes, angles of attack, and angles of sideslip to identify the conditions which produce the largest distortion and lowest pressure recovery. Pretest CFD was performed by NASA and Boeing, using multiple CFD codes, with various turbulence models. These data were used to make decisions regarding model integration, characterize inlet flow distortion patterns, and help define the wind tunnel test matrix. CFD was also performed post-test; when compared with test data, it was possible to make comparisons between measured model-scale and predicted full-scale distortion levels. This paper summarizes these CFD analyses.

ON THE PROBLEM OF CORRECTING TWISTED TURBINE BLADES,

DTIC Science & Technology

TURBINE BLADES , DESIGN), GAS TURBINES , STEAM TURBINES , BLADE AIRFOILS , ASPECT RATIO, FLUID DYNAMICS, SECONDARY FLOW, ANGLE OF ATTACK, INLET GUIDE VANES , CORRECTIONS, PERFORMANCE( ENGINEERING ), OPTIMIZATION, USSR

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.

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

An Experimental Investigation of Flow Conditions in the Vicinity of an NACA D(sub S)-type Cowling

NASA Technical Reports Server (NTRS)

Bryant, Rosemary P.; Boswinkle, Robert W.

1946-01-01

Data are presented of the flow conditions in the vicinity of an NACA D sub S -type cowling. Tests were made of a 1/2 scale-nacelle model at inlet-velocity ratios ranging from 0.23 to 1.02 and angles of attack from 6 deg to 10 deg. The velocity and direction of flow in the vertical plane of symmetry of the cowling were determined from orifices and tufts installed on a board aligned with the flow. Diagrams showing velocity ratio contours and lines of constant flow angles are given.

Secondary flow spanwise deviation model for the stators of NASA middle compressor stages

NASA Technical Reports Server (NTRS)

Roberts, W. B.; Sandercock, D. M.

1984-01-01

A model of the spanwise variation of deviation for stator blades is presented. Deviation is defined as the difference between the passage mean flow angle and the metal angle at the outlet of a blade element of an axial compressor stage. The variation of deviation is taken as the difference above or below that predicted by blade element, (i.e., two-dimensional) theory at any spanwise location. The variation of deviation is dependent upon the blade camber, solidity and inlet boundary layer thickness at the hub or tip end-wall, and the blade channel aspect ratio. If these parameters are known or can be calculated, the model provides a reasonable approximation of the spanwise variation of deviation for most compressor middle stage stators operating at subsonic inlet Mach numbers.

Axial and Centrifugal Compressor Mean Line Flow Analysis Method

NASA Technical Reports Server (NTRS)

Veres, Joseph P.

2009-01-01

This paper describes a method to estimate key aerodynamic parameters of single and multistage axial and centrifugal compressors. This mean-line compressor code COMDES provides the capability of sizing single and multistage compressors quickly during the conceptual design process. Based on the compressible fluid flow equations and the Euler equation, the code can estimate rotor inlet and exit blade angles when run in the design mode. The design point rotor efficiency and stator losses are inputs to the code, and are modeled at off design. When run in the off-design analysis mode, it can be used to generate performance maps based on simple models for losses due to rotor incidence and inlet guide vane reset angle. The code can provide an improved understanding of basic aerodynamic parameters such as diffusion factor, loading levels and incidence, when matching multistage compressor blade rows at design and at part-speed operation. Rotor loading levels and relative velocity ratio are correlated to the onset of compressor surge. NASA Stage 37 and the three-stage NASA 74-A axial compressors were analyzed and the results compared to test data. The code has been used to generate the performance map for the NASA 76-B three-stage axial compressor featuring variable geometry. The compressor stages were aerodynamically matched at off-design speeds by adjusting the variable inlet guide vane and variable stator geometry angles to control the rotor diffusion factor and incidence angles.

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.

Investigations of flowfields found in typical combustor geometries

NASA Technical Reports Server (NTRS)

Lilley, D. G.

1982-01-01

Experimental and theoretical research undertaken on 2-D axisymmetric geometries under low speed, nonreacting, turbulent, swirling flow conditions is reported. The flow enters the test section and proceeds into a larger chamber (the expansion ratio D/d = 2) via a sudden or gradual expansion (sidewall angle alpha = 90 and 45 degrees). Inlet swirl vanes are adjustable to a variety of vane angles with values of phi = 0, 38, 45, 60 and 70 degrees being emphasized.

Calculation of the flow field in supersonic mixed-compression inlets at angle of attack using the three-dimensional method of characteristics with discrete shock wave fitting

NASA Technical Reports Server (NTRS)

Vadyak, J.; Hoffman, J. D.

1978-01-01

The influence of molecular transport is included in the computation by treating viscous and thermal diffusion terms in the governing partial differential equations as correction terms in the method of characteristics scheme. The development of a production type computer program is reported which is capable of calculating the flow field in a variety of axisymmetric mixed-compression aircraft inlets. The results agreed well with those produced by the two-dimensional method characteristics when axisymmetric flow fields are computed. For three-dimensional flow fields, the results agree well with experimental data except in regions of high viscous interaction and boundary layer removal.

Numerical analysis of flow interaction of turbine system in two-stage turbocharger of internal combustion engine

NASA Astrophysics Data System (ADS)

Liu, Y. B.; Zhuge, W. L.; Zhang, Y. J.; Zhang, S. Y.

2016-05-01

To reach the goal of energy conservation and emission reduction, high intake pressure is needed to meet the demand of high power density and high EGR rate for internal combustion engine. Present power density of diesel engine has reached 90KW/L and intake pressure ratio needed is over 5. Two-stage turbocharging system is an effective way to realize high compression ratio. Because turbocharging system compression work derives from exhaust gas energy. Efficiency of exhaust gas energy influenced by design and matching of turbine system is important to performance of high supercharging engine. Conventional turbine system is assembled by single-stage turbocharger turbines and turbine matching is based on turbine MAP measured on test rig. Flow between turbine system is assumed uniform and value of outlet physical quantities of turbine are regarded as the same as ambient value. However, there are three-dimension flow field distortion and outlet physical quantities value change which will influence performance of turbine system as were demonstrated by some studies. For engine equipped with two-stage turbocharging system, optimization of turbine system design will increase efficiency of exhaust gas energy and thereby increase engine power density. However flow interaction of turbine system will change flow in turbine and influence turbine performance. To recognize the interaction characteristics between high pressure turbine and low pressure turbine, flow in turbine system is modeled and simulated numerically. The calculation results suggested that static pressure field at inlet to low pressure turbine increases back pressure of high pressure turbine, however efficiency of high pressure turbine changes little; distorted velocity field at outlet to high pressure turbine results in swirl at inlet to low pressure turbine. Clockwise swirl results in large negative angle of attack at inlet to rotor which causes flow loss in turbine impeller passages and decreases turbine efficiency. However negative angle of attack decreases when inlet swirl is anti-clockwise and efficiency of low pressure turbine can be increased by 3% compared to inlet condition of clockwise swirl. Consequently flow simulation and analysis are able to aid in figuring out interaction mechanism of turbine system and optimizing turbine system design.

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

Effect of reducing rotor blade inlet diameter on the performance of a 11.66-Centimeter radial-inflow turbine

NASA Technical Reports Server (NTRS)

Kofskey, M. G.; Haas, J. E.

1973-01-01

The effect of increased rotor blade loading on turbine performance was investigated by reducing rotor blade inlet diameter. The reduction was made in four stages. Each modification was tested with the same stator using cold air as the working fluid. Results are presented in terms of equivalent mass flow and efficiency at equivalent design rotative speed and over a range of pressure ratios. Internal flow characteristics are shown in terms of stator exit static pressure and the radial variation of local loss and rotor-exit flow angle with radius ratio. Included are velocity diagrams calculated from the experimental results.

Flow visualization of lateral jet injection into swirling crossflow

NASA Technical Reports Server (NTRS)

Ferrell, G. B.; Aoki, K.; Lilley, D. G.

1985-01-01

Flow visualization experiments have been conducted to characterize the time-mean flowfield of a deflected turbulent jet in a confining cylindrical crossflow. Jet-to-crossflow velocity ratios of 2, 4, and 6 were investigated, under crossflow inlet swirler vane angles of 0 (swirler removed), 45 and 70 degrees. Smoke, neutrally-buoyant helium-filled soap bubbles, and multi-spark flow visualization were employed to highlight interesting features of the deflected jet, as well as the trajectory and spread pattern of the jet. Gross flowfield characterization was obtained for a range of lateral jet-to-crossflow velocity ratios and a range of inlet swirl strengths in the main flow. The flow visualization results agree well with the measurements obtained elsewhere with the six-orientation single hot-wire method.

Management of Total Pressure Recovery, Distortion and High Cycle Fatigue in Compact Air Vehicle Inlets

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Baust, Henry D.; Agrell, Johan

2002-01-01

It is the purpose of this study to demonstrate the viability and economy of Response Surface Methods (RSM) and Robustness Design Concepts (RDC) to arrive at micro-secondary flow control installation designs that maintain optimal inlet performance over a range of the mission variables. These statistical design concepts were used to investigate the robustness properties of 'low unit strength' micro-effector installations. '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.

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.

Evaluation of a hemispherical head flow direction sensor for inlet duct measurements

NASA Technical Reports Server (NTRS)

Bennett, D. L.

1975-01-01

A hemispherical head flow direction sensor was tested in a wind tunnel to evaluate its effectiveness for measuring dynamic duct flow direction angles of plus and minus 27 degrees. The tests were conducted at Reynolds numbers of 3.8 million per meter (1.0 million per foot) and 4.92 million per meter (1.5 million per foot) and at Mach numbers from 0.30 to 0.70. The design criteria for the probe are discussed and the wind tunnel results are presented. Three techniques for deriving the flow angles are described.

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

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.

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.

Steady and Unsteady Velocity Measurements in a Small Turbocharger Turbine with Computational Validation

NASA Astrophysics Data System (ADS)

Karamanis, N.; Palfreyman, D.; Arcoumanis, C.; Martinez-Botas, R. F.

2006-07-01

The detailed flow characteristics of three high-pressure-ratio mixed-flow turbines were investigated under both steady and pulsating flow conditions. Two rotors featured a constant inlet blade angle, one with 12 blades and the second with 10. The third rotor was shorter and had a nominally constant incidence angle. The rotors find application on an automotive high-speed large commercial diesel turbocharger. The steady flow entering and exiting the blades has been quantified by a laser Doppler velocimetry system. The measurements were performed at a plane 3.0-mm ahead of the rotor leading edge and 9.5-mm downstream the rotor trailing edge. The turbine test conditions corresponded to the peak efficiency point at two rotational speeds, 29,400 and 41,300-rpm. The results were resolved in a blade-to-blade sense to examine fully the nature of the flow at turbocharger representative conditions. A correlation between the combined effects of incidence and exit flow angle with the isentropic efficiency has been verified. Regarding pulsating flow, the velocity data and their corresponding instantaneous velocity triangles were resolved in a blade-to-blade sense to understand better the complex phenomenon. The results highlighted the potential of a nominally constant incidence design to absorb better the inadequacy of the volute to discharge the exhaust gas uniformly along the blade leading edge. A double vortex rotating in a clockwise sense propagated on the plane normal to the meridional direction. This should be attributed to the effect of the passing blade that was acting as a blockage to the flow. The phenomenon was more pronounced near the suction and pressure surfaces of the blade, but diminished at the mid-passage region where the flow exhibited its best level of guidance. The full mixed flow turbine stage under transient conditions was modelled firstly with a 'steady' inlet and secondly with a 'pulsating' inlet boundary condition. In both cases comparison was made to experiment performance and LDV measurements. With the steady inlet boundary condition, a high level of accuracy was achieved when compared to the experimental performance and velocity field. The velocity along the leading edge showed the same discrepancy as the single passage analysis that is with the radial and axial component from mid span to the blade tip. At the trailing edge features identified in the experimental data are identified in the numerical results; the velocity field appears more 'diffused' across the plane as per the experimental data than from the single passage analysis. With the pulsating inlet boundary, the predicted velocity traces in the volute and close to the turbine lead and trailing edge show excellent agreement in both form (against time) and magnitude.

Vortex nozzle for segmenting and transporting metal chips from turning operations

DOEpatents

Bieg, L.F.

1993-04-20

Apparatus for collecting, segmenting and conveying metal chips from machining operations utilizes a compressed gas driven vortex nozzle for receiving the chip and twisting it to cause the chip to segment through the application of torsional forces to the chip. The vortex nozzle is open ended and generally tubular in shape with a converging inlet end, a constant diameter throat section and a diverging exhaust end. Compressed gas is discharged through angled vortex ports in the nozzle throat section to create vortex flow in the nozzle and through an annular inlet at the entrance to the converging inlet end to create suction at the nozzle inlet and cause ambient air to enter the nozzle. The vortex flow in the nozzle causes the metal chip to segment and the segments thus formed to pass out of the discharge end of the nozzle where they are collected, cleaned and compacted as needed.

Installed F/A-18 inlet flow calculations at 30 degrees angle-of-attack: A comparative study

NASA Technical Reports Server (NTRS)

Smith, C. Frederic; Podleski, Steve D.

1994-01-01

NASA Lewis is currently engaged in a research effort as a team member of the High Alpha Technology Program (HATP) within NASA. This program utilizes a specially equipped F/A-18, the High Alpha Research Vehicle (HARV), in an ambitious effort to improve the maneuverability of high-performance military aircraft at low subsonic speed, high angle of attack conditions. The overall objective of the Lewis effort is to develop inlet technology that will ensure efficient airflow delivery to the engine during these maneuvers. One part of the Lewis approach utilizes computational fluid dynamics codes to predict the installed performance of inlets for these highly maneuverable aircraft. Full Navier-Stokes (FNS) calculations on the installed F/A-18 inlet at 30 degrees angle of attack, 0 degrees yaw, and a freestream Mach number of 0.2 have been obtained in this study using an algebraic turbulence model with two grids (original and revised). Results obtained with the original grid were used to determine where further grid refinements and additional geometry were needed. In order to account properly for the external effects, the forebody, leading edge extension (LEX), ramp, and wing were included with inlet geometry. In the original grid, the diverter, LEX slot, and leading edge flap were not included due to insufficient geometry definition, but were included in a revised grid. In addition, a thin-layer Navier-Stokes (TLNS) code is used with the revised grid and the numerical results are compared to those obtained with the FNS code. The TLNS code was used to evaluate the effects on the solution using a code with more recent CFD developments such as upwinding with TVD schemes versus central differencing with artificial dissipation. The calculations are compared to a limited amount of available experimental data. The predicted forebody/fuselage surface static pressures compared well with data of all solutions. The predicted trajectory of the vortex generated under the LEX was different for each solution. These discrepancies are attributed to differences in the grid resolution and turbulence modeling. All solutions predict that this vortex is ingested by the inlet. The predicted inlet total pressure recoveries are lower than data and the distortions are higher than data. The results obtained with the revised grid were significantly improved from the original grid results. The original grid results indicated the ingested vortex migrated to the engine face and caused additional distortions to those already present due to secondary flow development. The revised grid results indicate that the ingested vortex is dissipated along the inlet duct inboard wall. The TLNS results indicate the flow at the engine face was much more distorted than the FNS results and is attributed to the pole boundary condition introducing numerical distortions into the flow field.

Aerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade at Low Inlet Turbulence

NASA Technical Reports Server (NTRS)

Flegel-McVetta, Ashlie B.; Giel, Paul W.; Welch, Gerard E.

2013-01-01

Aerodynamic measurements obtained in a transonic linear cascade were used to assess the impact of large incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Steady-state data were obtained for ten incidence angles ranging from +15.8 deg to -51.0 deg. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12×10(exp 5) to 2.12×10(exp 6). Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8 deg and -36.7 deg. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures. Tests were conducted in the NASA Glenn Transonic Turbine Blade Cascade Facility. The measurements reflect strong secondary flows associated with the high aerodynamic loading levels at large positive incidence angles and an increase in loss levels with decreasing Reynolds number. The secondary flows decrease with negative incidence as the blade becomes unloaded. Transitional flow is admitted in this low inlet turbulence dataset, making it a challenging CFD test case. The dataset will be used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range. deg

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

Study on the effects of flow in the volute casing on the performance of a sirocco fan

NASA Astrophysics Data System (ADS)

Adachi, Tsutomu; Sugita, Naohiro; Ohomori, Satoshi

2004-08-01

The flow at the exit from the runner blade of a centrifugal fan with forward curved blades (a sirocco fan) sometimes separates and becomes unstable. We have conducted many researches on the impeller shape of a sirocco fan, proper inlet and exit blade angles were considered to obtain optimum performance. In this paper, the casing shape were decided by changing the circumferential angle, magnifying angle and the width, 21 sorts of casings were used. Performance tests, inner flow velocity and pressure distributions were measured as well. Computational fluid dynamic calculations were also made and compared with the experimental results. Finally, the most suitable casing shape for best performance is considered.

Forward velocity effects on fan noise and the suppression characteristics of advanced inlets as measured in the NASA-Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Moore, M. T.

1980-01-01

Forward velocity effects on the forward radiated fan noise and on the suppression characteristics of three advanced inlets relative to a baseline cylindrical inlet were measured in the NASA Ames Research Center 40 x 80 foot Wind Tunnel. A modified JT15D turbofan engine in a quiet nacelle was the source of fan noise; the advanced inlets were a Conventional Takeoff/Landing (CTOL) hybrid inlet, a Short Takeoff/Landing (STOL) hybrid inlet, and a treated deflector inlet. Also measured were the static to flight effects on the fan noise of canting the baseline inlet 4 deg downward to simulate typical wing mounted turbofan engines. The CTOL hybrid inlet suppressed the high tip speed fan noise as much as 18 PNdB on a 61 m (200 ft) sideline scaled to a CF6 size engine while the STOL hybrid inlet suppressed the low tip speed fan noise as much as 13 PNdB on a 61 m (200 ft) sideline scaled to a OCSEE size engine. The deflector inlet suppressed the high tip speed fan noise as much as 13 PNdB at 61 m (200 ft) overhead scaled to a CF6 size engine. No significant changes in fan noise suppression for the CTOL and STOL hybrid inlets occurred for forward velocity changes above 21 m/s (68 ft/s) or for angle of attack changes up to 15 deg. However, changes in both forward velocity and angle of attack changed the deflector inlet noise unpredictably due to the asymmetry of the inlet flow field into the fan.

Exergy Analyses of Fabricated Compound Parabolic Solar Collector with Evacuated Tubes at Different Operating Conditions: Indore (India)

NASA Astrophysics Data System (ADS)

Geete, Ankur; Dubey, Akash; Sharma, Ankush; Dubey, Anshul

2018-05-01

In this research work, compound parabolic solar collector (CPC) with evacuated tubes is fabricated. Main benefit of CPC is that there is no requirement of solar tracking system. With fabricated CPC; outlet temperatures of flowing fluid, instantaneous efficiencies, useful heat gain rates and inlet exergies (with and without considering Sun's cone angle) are experimentally found. Observations are taken at different time intervals (1200, 1230, 1300, 1330 and 1400 h), mass flow rates (1.15, 0.78, 0.76, 0.86 and 0.89 g/s), ambient temperatures and with various dimensions of solar collector. This research work is concluded as; maximum instantaneous efficiency is 69.87% which was obtained with 0.76 g/s flow rate of water at 1300 h and 42°C is the maximum temperature difference which was also found at same time. Maximum inlet exergies are 139.733 and 139.532 kW with and without considering Sun's cone angle at 1300 h, respectively. Best thermal performance from the fabricated CPC with evacuated tubes is found at 1300 h. Maximum inlet exergy is 141.365 kW which was found at 1300 h with 0.31 m aperture width and 1.72 m absorber pipe length.

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.

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.

Experimental investigation of a 0.15 scale model of a conformal variable-ramp inlet for the F-16 airplane

NASA Technical Reports Server (NTRS)

Hawkins, J. E.

1980-01-01

A 0.15 scale model of a proposed conformal variable-ramp inlet for the Multirole Fighter was tested from Mach 0.8 to 2.2 at a wide range of angles of attack and sideslip. Inlet ramp angle was varied to optimize ramp angle as a function of engine airflow, Mach number, angle of attack, and angle of sideslip. Several inlet configuration options were investigated to study their effects on inlet operation and to establish the final flight configuration. These variations were cowl sidewall cutback, cowl lip bluntness, boundary layer bleed, and first-ramp leading edge shape. Diagnostic and engine face instrumentation were used to evaluate inlet operation at various inlet stations and at the inlet/engine interface. Pressure recovery and stability of the inlet were satisfactory for the proposed application. On the basis of an engine stability audit of the worst-case instantaneous distortion patterns, no inlet/engine compatibility problems are expected for normal operations.

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.

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

Bu, L.; Zhao, J.

The supercritical water mixing phenomenon is investigated with a wide range of conditions, i.e. the inlet temperature of the streams ranges from 323.15 K to 723.15 K and the pressure ranges from 25 MPa to 45 MPa. A sensitivity study is carried out for the jet and main flow velocity ratio (VR) which is varying from 1 to 40. In addition, the effect of the inject angles of branch flow to main flow on the mixing is conducted by varying the inject angle from 80 deg. to 100 deg.. The results show that the maximum temperature gradient appears on themore » wall of the upstream side in all the cases, and the inclined angles can be optimized to mitigate the thermal stress. (authors)« less

Pressure distributions for a rectangular supersonic inlet at subsonic speeds

NASA Technical Reports Server (NTRS)

Fuller, D. E.

1976-01-01

Pressure distribution data are provided for a supersonic rectangular inlet at subsonic speeds. Variations in cowl and ramp geometry as well as sideplate sweep were investigated. Tests were made in the Langley 16-foot transonic tunnel and the Langley high speed 7- by 10-foot tunnel for Mach numbers of 0.6, 0.7, and 0.8. Angles of attack investigated were 0 deg, 4 deg, and 8 deg for a range of mass flow ratios.

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.

Analytical correlation of centrifugal compressor design geometry for maximum efficiency with specific speed

NASA Technical Reports Server (NTRS)

Galvas, M. R.

1972-01-01

Centrifugal compressor performance was examined analytically to determine optimum geometry for various applications as characterized by specific speed. Seven specific losses were calculated for various combinations of inlet tip-exit diameter ratio, inlet hub-tip diameter ratio, blade exit backsweep, and inlet-tip absolute tangential velocity for solid body prewhirl. The losses considered were inlet guide vane loss, blade loading loss, skin friction loss, recirculation loss, disk friction loss, vaneless diffuser loss, and vaned diffuser loss. Maximum total efficiencies ranged from 0.497 to 0.868 for a specific speed range of 0.257 to 1.346. Curves of rotor exit absolute flow angle, inlet tip-exit diameter ratio, inlet hub-tip diameter ratio, head coefficient and blade exit backsweep are presented over a range of specific speeds for various inducer tip speeds to permit rapid selection of optimum compressor size and shape for a variety of applications.

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.

Laser measurements of unsteady flow field in a radial turbine guide vanes

NASA Astrophysics Data System (ADS)

Pasin, M.; Tabakoff, W.

1992-01-01

Detailed measurements of a unsteady flow field within the inlet guide vanes (IGV) of a radial inflow turbine were performed using a three component Laser Doppler Velocimeter (LDV) system. The mean velocity, the flow angle and the turbulence results are presented at the midspan plane for different rotor positions. These results are compared with the measurements obtained in the same passage in the absence of the rotor.

The effects of design and operating variables on the response of an axial flow fan to inlet flow distortions. M.S. Thesis

NASA Technical Reports Server (NTRS)

Yocum, A. M., II

1978-01-01

The results of a study of total pressure and velocity circumferential distortions in an axial-flow fan are presented. Some of the fundamental experimental data needed to understand distorted flow phenomena as affected by design and operating variables are provided. The flow through an isolated rotor was examined at various operating conditions with six different distortions and three different blade stagger angles. Circumferential surveys were conducted upstream and downstream of the rotor using five-hole probes in the nonnulling mode. The total pressure and axial velocity distortion data were analyzed to determine the degree of distortion attenuation as a function of blade stagger angle, mean incidence angle, and reduced frequency. The results indicate that, for the rotors tested, the mean incidence or loading has very little effect on the distortion attenuation.

Tactical missile aerodynamics - General topics. Progress in Astronautics and Aeronautics. Vol. 141

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

Hemsch, M.J.

1992-01-01

The present volume discusses the development history of tactical missile airframes, aerodynamic considerations for autopilot design, a systematic method for tactical missile design, the character and reduction of missile observability by radar, the visualization of high angle-of-attack flow phenomena, and the behavior of low aspect ratio wings at high angles of attack. Also discussed are airbreathing missile inlets, 'waverider' missile configurations, bodies with noncircular cross-sections and bank-to-turn missiles, asymmetric flow separation and vortex shedding on bodies-of-revolution, unsteady missile flows, swept shock-wave/boundary-layer interactions, pylon carriage and separation of stores, and internal stores carriage and separation.

STGSTK: A computer code for predicting multistage axial flow compressor performance by a meanline stage stacking method

NASA Technical Reports Server (NTRS)

Steinke, R. J.

1982-01-01

A FORTRAN computer code is presented for off-design performance prediction of axial-flow compressors. Stage and compressor performance is obtained by a stage-stacking method that uses representative velocity diagrams at rotor inlet and outlet meanline radii. The code has options for: (1) direct user input or calculation of nondimensional stage characteristics; (2) adjustment of stage characteristics for off-design speed and blade setting angle; (3) adjustment of rotor deviation angle for off-design conditions; and (4) SI or U.S. customary units. Correlations from experimental data are used to model real flow conditions. Calculations are compared with experimental data.

Design of 9.271-pressure-ratio 5-stage core compressor and overall performance for first 3 stages

NASA Technical Reports Server (NTRS)

Steinke, Ronald J.

1986-01-01

Overall aerodynamic design information is given for all five stages of an axial flow core compressor (74A) having a 9.271 pressure ratio and 29.710 kg/sec flow. For the inlet stage group (first three stages), detailed blade element design information and experimental overall performance are given. At rotor 1 inlet tip speed was 430.291 m/sec, and hub to tip radius ratio was 0.488. A low number of blades per row was achieved by the use of low-aspect-ratio blading of moderate solidity. The high reaction stages have about equal energy addition. Radial energy varied to give constant total pressure at the rotor exit. The blade element profile and shock losses and the incidence and deviation angles were based on relevant experimental data. Blade shapes are mostly double circular arc. Analysis by a three-dimensional Euler code verified the experimentally measured high flow at design speed and IGV-stator setting angles. An optimization code gave an optimal IGV-stator reset schedule for higher measured efficiency at all speeds.

Investigations of the influence of the profile thickness of the compressible plane flow through compressor cascades

NASA Technical Reports Server (NTRS)

Bahr, J.

1978-01-01

Flow-through cascade of an aircraft turbine compressor is studied experimentally over wide range of Reynolds numbers and subsonic Mach numbers; it was found that deterioration of flow properties due to decreasing Reynolds numbers is less noticeable on thin profiles than on thick ones; however, thick profiles are advantageous in compressors designed for efficient partial load behavior because thick profiles have a relatively large range of usable inlet flow angles.

Parametric Study of Sealant Nozzle

NASA Astrophysics Data System (ADS)

Yamamoto, Yoshimi

It has become apparent in recent years the advancement of manufacturing processes in the aerospace industry. Sealant nozzles are a critical device in the use of fuel tank applications for optimal bonds and for ground service support and repair. Sealants has always been a challenging area for optimizing and understanding the flow patterns. A parametric study was conducted to better understand geometric effects of sealant flow and to determine whether the sealant rheology can be numerically modeled. The Star-CCM+ software was used to successfully develop the parametric model, material model, physics continua, and simulate the fluid flow for the sealant nozzle. The simulation results of Semco sealant nozzles showed the geometric effects of fluid flow patterns and the influences from conical area reduction, tip length, inlet diameter, and tip angle parameters. A smaller outlet diameter induced maximum outlet velocity at the exit, and contributed to a high pressure drop. The conical area reduction, tip angle and inlet diameter contributed most to viscosity variation phenomenon. Developing and simulating 2 different flow models (Segregated Flow and Viscous Flow) proved that both can be used to obtain comparable velocity and pressure drop results, however; differences are seen visually in the non-uniformity of the velocity and viscosity fields for the Viscous Flow Model (VFM). A comprehensive simulation setup for sealant nozzles was developed so other analysts can utilize the data.

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.

Inlet Distortion for an F/A-18A Aircraft During Steady Aerodynamic Conditions up to 60 deg Angle of Attack

NASA Technical Reports Server (NTRS)

Walsh, Kevin R.; Yuhas, Andrew J.; Williams, John G.; Steenken, William G.

1997-01-01

The effects of high-angle-of-attack flight on aircraft inlet aerodynamic characteristics were investigated at NASA Dryden Flight Research Center, Edwards, California, as part of NASA's High Alpha Technology Program. The highly instrumented F/A-18A High Alpha Research Vehicle was used for this research. A newly designed inlet total-pressure rake was installed in front of the starboard F404-GE-400 engine to measure inlet recovery and distortion characteristics. One objective was to determine inlet total-pressure characteristics at steady high-angle-of-attack conditions. Other objectives include assessing whether significant differences exist in inlet distortion between rapid angle-of-attack maneuvers and corresponding steady aerodynamic conditions, assessing inlet characteristics during aircraft departures, providing data for developing and verifying computational fluid dynamic codes, and calculating engine airflow using five methods. This paper addresses the first objective by summarizing results of 79 flight maneuvers at steady aerodynamic conditions, ranging from -10 deg to 60 deg angle of attack and from -8 deg to 11 deg angle of sideslip at Mach 0.3 and 0.4. These data and the associated database have been rigorously validated to establish a foundation for understanding inlet characteristics at high angle of attack.

Computational study of fuel injection in a shcramjet inlet

NASA Astrophysics Data System (ADS)

Parent, Bernard

The primary objective of this investigation is to present the mixing of fuel with air in the inlet of a shock-induced combustion ramjet (shcramjet). The study is limited to non-reacting hydrogen-air mixing in an external-compression inlet at a flight Mach number of 11 and at a dynamic pressure of 1400 psf (67032 Pa), using an array of cantilevered ramp injectors. A numerical method based on the Yee-Roe scheme and block-implicit approximate factorization is developed to solve the FANS equations closed by the Wilcox ko turbulence model. A new acceleration technique for streamwise-separated hypersonic flow, dubbed the "marching window", is presented. The dilatational dissipation correction is seen to affect the mixing efficiency considerably for a cantilevered ramp injector flowfield even at a vanishing convective Mach number, due to the high turbulent Mach number generated by the high cross-stream shear induced by the ramp-generated axial vortices. Due to the fuel being injected at a very high speed, fuel injection in the inlet is found to increase considerably the thrust potential, with a gain exceeding the loss by 40--120%. Losses due to skin friction are seen to play a significant role in the inlet, as they are estimated to make up as much as 50--70% of the thrust potential losses. The use of a turbulence model that can predict accurately the wall shear stress is hence crucial in assessing the losses accurately in a shcramjet inlet. Substituting the second inlet shock by a Prandtl-Meyer compression fan is encouraged as it decreases the thrust potential losses, reduces the risk of premature ignition by reducing the static temperature, while decreasing the mixing efficiency by a mere 6%. One approach that is observed herein to be successful at increasing the mixing efficiency in the inlet is by alternating the injection angle along the injector array. The use of two injection angles of 9 and 16 degrees is seen to result in a 32% increase in the mixing efficiency at the expense of a 14% increase in the losses when compared to a single injection angle of 10 degrees. Using alternating injection angles, the mixing efficiency reaches as much as 0.47 at the inlet exit.

Engine inlet distortion in a 9.2 percent scaled vectored thrust STOVL model in ground effect

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Neiner, George; Flood, J. D.; Amuedo, K. C.; Strock, T. W.

1989-01-01

Advanced Short Takeoff/Vertical Landing (STOVL) aircraft which can operate from remote locations, damaged runways, and small air capable ships are being pursued for deployment around the turn of the century. To achieve this goal, a cooperative program has been defined for testing in the NASA Lewis 9- by 15-foot Low Speed Wind Tunnel (LSWT) to establish a database for hot gas ingestion, one of the technologies critical to STOVL. This paper presents results showing the engine inlet distortions (both temperature and pressure) in a 9.2 percent scale Vectored Thrust STOVL model in ground effects. Results are shown for the forward nozzle splay angles of 0, -6, and 18 deg. The model support system had 4 deg of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity was varied from 8 to 23 kn.

Aerodynamic Measurements of an Incidence Tolerant Blade in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

McVetta, Ashlie B.; Giel, Paul W.

2012-01-01

An overview of the recent facility modifications to NASA s Transonic Turbine Blade Cascade Facility and aerodynamic measurements on the VSPT incidence-tolerant blade are presented. This work supports the development of variable-speed power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50% speed range from takeoff to altitude cruise. This results in 50 or more variations in VSPT blade incidence angles. The Transonic Turbine Blade Cascade Facility has the ability to operate over a wide range of Reynolds numbers and Mach numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation. Details of the modifications are described. An incidence-tolerant blade was developed under an RTPAS study contract and tested in the cascade to look at the effects of large incidence angle and Reynolds number variations. Recent test results are presented which include midspan exit total pressure and flow angle measurements obtained at three inlet angles representing the cruise, take-off, and maximum incidence flight mission points. For each inlet angle, data were obtained at five flow conditions with exit Reynolds numbers varying from 2.12 106 to 2.12 105 and two isentropic exit Mach numbers of 0.72 and 0.35. Three-dimensional flowfield measurements were also acquired at the cruise and take-off points. The flowfield measurements were acquired using a five-hole and three-hole pneumatic probe located in a survey plane 8.6% axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure distributions were also acquired for each flow condition.

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.

Experimental Evaluation of Inlet Distortion on an Ejector Powered Hybrid Wing Body at Take-off and Landing Conditions

NASA Technical Reports Server (NTRS)

Carter, Melissa B.; Shea, Patrick R.; Flamm, Jeffrey D.; Schuh, Michael; James, Kevin D.; Sexton, Matthew R.; Tompkins, Daniel M.; Beyar, Michael D.

2016-01-01

As part of the NASA Environmentally Responsible Aircraft project, an ultra high bypass ratio engine integration on a hybrid wing body demonstration was planned. The goal was to include engine and airframe integration concepts that reduced fuel consumption by at least 50% while still reducing noise 42 db cumulative on the ground. Since the engines would be mounted on the upper surface of the aft body of the aircraft, the inlets may be susceptible to vortex ingestion from the wing leading edge at high angles of attack and sideslip, and separated wing/body flow. Consequently, experimental and computational studies were conducted to collect flow surveys useful for characterizing engine operability. The wind tunnel tests were conducted at two NASA facilities, the 14- by 22-foot at NASA Langley and the 40- by 80-foot at NASA Ames Research Center. The test results included in this paper show that the distortion and pressure recovery levels were acceptable for engine operability. The CFD studies conducted to compare to experimental data showed excellent agreement for the angle of attacks examined, although failed to match the low speed experimental data at high sideslip angles.

Experimental Investigation of Two-Phase Oil (D130)-Water Flow in 4″ Pipe for Different Inclination Angles

NASA Astrophysics Data System (ADS)

Shaahid, S. M.; Basha, Mehaboob; Al-Hems, Luai M.

2018-03-01

Oil and water are often produced and transported together in pipelines that have various degrees of inclination from the horizontal. The flow of two immiscible liquids oil and water in pipes has been a research topic since several decades. In oil and chemical industries, knowledge of the frictional pressure loss in oil-water flows in pipes is necessary to specify the size of the pump required to pump the emulsions. An experimental investigation has been carried out for measurement of pressure drop of oil (D130)-water two-phase flows in 4 inch diameter inclined stainless steel pipe at different flow conditions. Experiments were conducted for different inclination angles including; 0°, 15°, 30° (for water cuts “WC” 0 - 100%). The flow rates at the inlet were varied from 4000 to 8000 barrels-per-day (BPD). For a given flow rate the frictional pressure drop has been found to increase (for all angles) from WC = 0 - 60%, and thereafter friction pressure drop decreases, this could be due phase inversion. For a given WC 40%, the frictional pressure drop has been found to increase with angle and flow rate. It has been noticed that inclination angle has appreciable effect on frictional pressure drop.

Dynamic interactions between hypersonic vehicle aerodynamics and propulsion system performance

NASA Technical Reports Server (NTRS)

Flandro, G. A.; Roach, R. L.; Buschek, H.

1992-01-01

Described here is the development of a flexible simulation model for scramjet hypersonic propulsion systems. The primary goal is determination of sensitivity of the thrust vector and other system parameters to angle of attack changes of the vehicle. Such information is crucial in design and analysis of control system performance for hypersonic vehicles. The code is also intended to be a key element in carrying out dynamic interaction studies involving the influence of vehicle vibrations on propulsion system/control system coupling and flight stability. Simple models are employed to represent the various processes comprising the propulsion system. A method of characteristics (MOC) approach is used to solve the forebody and external nozzle flow fields. This results in a very fast computational algorithm capable of carrying out the vast number of simulation computations needed in guidance, stability, and control studies. The three-dimensional fore- and aft body (nozzle) geometry is characterized by the centerline profiles as represented by a series of coordinate points and body cross-section curvature. The engine module geometry is represented by an adjustable vertical grid to accommodate variations of the field parameters throughout the inlet and combustor. The scramjet inlet is modeled as a two-dimensional supersonic flow containing adjustable sidewall wedges and multiple fuel injection struts. The inlet geometry including the sidewall wedge angles, the number of injection struts, their sweepback relative to the vehicle reference line, and strut cross-section are user selectable. Combustion is currently represented by a Rayleigh line calculation including corrections for variable gas properties; improved models are being developed for this important element of the propulsion flow field. The program generates (1) variation of thrust magnitude and direction with angle of attack, (2) pitching moment and line of action of the thrust vector, (3) pressure and temperature distributions throughout the system, and (4) performance parameters such as thrust coefficient, specific impulse, mass flow rates, and equivalence ratio. Preliminary results are in good agreement with available performance data for systems resembling the NASP vehicle configuration.

Optimization study for high speed radial turbine with special reference to design variables

NASA Technical Reports Server (NTRS)

Khalil, I.; Tabakoff, W.

1977-01-01

Numerical results of a theoretical investigation are presented to provide information about the effect of variation of the different design and operating parameters on radial inflow turbine performance. The effects of variations in the mass flow rate, rotor tip Mach number, inlet flow angles, number of rotor blades and hub to shroud radius ratio, on the internal fluid dynamics of turbine rotors, was investigated. A procedure to estimate the flow deviation angles at the turbine exit is also presented and used to examine the influence of the operating conditions and the rotor geometrical configuration on these deviations. The significance of the results obtained is discussed with respect to improved turbine performance.

Optical/Infrared Properties of Atmospheric Aerosols with an In-Situ, Multi-Wavelength, Multichannel Nephelometer System.

DTIC Science & Technology

1985-04-01

from any angle of approach. An angular impetus is imparted to the particle 25 7. motion via the eight evenly spaced entrance vanes. As the particles...measurement cycle. 2 f.. 27 VA NE S M= VANE ASSEMBLY BASE IN SECT SCREEN TUBE PROTECTIVE CPLUGHOUSING HOUSING-DEFLECTORAEOYMIINT SPACING PHAEO YNAI N...AERODYNAMIC FLOW DEFLECTOR OUTER TUJBE TAPERED p. Figure 3 r Wedding PM 1 0 Inlet [I]J 28 . .I .. II | I I J . 1 _1 ! I 7 I ! . ?. . . - 120 -- WEDDING INLET

Cross spectra between temperature and pressure in a constant area duct downstream of a combustor

NASA Technical Reports Server (NTRS)

Miles, J. H.; Wasserbauer, C. A.; Krejsa, E. A.

1983-01-01

The feasibility of measuring pressure temperature cross spectra and coherence and temperature-temperature cross spectra and coherence at spatially separated points along with pressure and temperature auto-spectra in a combustion rig was investigated. The measurements were made near the inlet and exit of a 6.44 m long duct attached to a J-47 combustor. The fuel used was Jet A. The cross spectra and coherence measurements show the pressure and temperature fluctuations correlate best at low frequencies. At the inlet the phenomena controlling the phase relationship between pressure and temperature could not be identified. However, at the duct exit the phase angle of the pressure is related to the phase angle of the temperature by the convected flow time delay.

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.

Three-Dimensional Flow Field Measurements in a Transonic Turbine Cascade

NASA Technical Reports Server (NTRS)

Giel, P. W.; Thurman, D. R.; Lopez, I.; Boyle, R. J.; VanFossen, G. J.; Jett, T. A.; Camperchioli, W. P.; La, H.

1996-01-01

Three-dimensional flow field measurements are presented for a large scale transonic turbine blade cascade. Flow field total pressures and pitch and yaw flow angles were measured at an inlet Reynolds number of 1.0 x 10(exp 6) and at an isentropic exit Mach number of 1.3 in a low turbulence environment. Flow field data was obtained on five pitchwise/spanwise measurement planes, two upstream and three downstream of the cascade, each covering three blade pitches. Three-hole boundary layer probes and five-hole pitch/yaw probes were used to obtain data at over 1200 locations in each of the measurement planes. Blade and endwall static pressures were also measured at an inlet Reynolds number of 0.5 x 10(exp 6) and at an isentropic exit Mach number of 1.0. Tests were conducted in a linear cascade at the NASA Lewis Transonic Turbine Blade Cascade Facility. The test article was a turbine rotor with 136 deg of turning and an axial chord of 12.7 cm. The flow field in the cascade is highly three-dimensional as a result of thick boundary layers at the test section inlet and because of the high degree of flow turning. The large scale allowed for very detailed measurements of both flow field and surface phenomena. The intent of the work is to provide benchmark quality data for CFD code and model verification.

Inflow/outflow boundary conditions for particle-based blood flow simulations: Application to arterial bifurcations and trees

DOE PAGES

Lykov, Kirill; Li, Xuejin; Lei, Huan; ...

2015-08-28

When blood flows through a bifurcation, red blood cells (RBCs) travel into side branches at different hematocrit levels, and it is even possible that all RBCs enter into one branch only, leading to a complete separation of plasma and R- BCs. To quantify this phenomenon via particle-based mesoscopic simulations, we developed a general framework for open boundary conditions in multiphase flows that is effective even for high hematocrit levels. The inflow at the inlet is duplicated from a fully developed flow generated in a pilot simulation with periodic boundary conditions. The outflow is controlled by adaptive forces to maintain themore » flow rate and velocity gradient at fixed values, while the particles leaving the arteriole at the outlet are removed from the system. Upon valida- tion of this approach, we performed systematic 3D simulations to study plasma skimming in arterioles of diameters 20 to 32 microns. For a flow rate ratio 6:1 at the branches, we observed the \\all-or-nothing" phenomenon with plasma only entering the low flow rate branch. We then simulated blood-plasma separation in arteriolar bifurcations with different bifurcation angles and same diameter of the daughter branches. Our simulations predict a significant increase in RBC flux through the main daughter branch as the bifurcation angle is increased. Lastly, we demonstrated the new methodology for simulating blood flow in ves- sels with multiple inlets and outlets, constructed using an angiogenesis model.« less

Inflow/outflow boundary conditions for particle-based blood flow simulations: Application to arterial bifurcations and trees

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

Lykov, Kirill; Li, Xuejin; Lei, Huan

When blood flows through a bifurcation, red blood cells (RBCs) travel into side branches at different hematocrit levels, and it is even possible that all RBCs enter into one branch only, leading to a complete separation of plasma and R- BCs. To quantify this phenomenon via particle-based mesoscopic simulations, we developed a general framework for open boundary conditions in multiphase flows that is effective even for high hematocrit levels. The inflow at the inlet is duplicated from a fully developed flow generated in a pilot simulation with periodic boundary conditions. The outflow is controlled by adaptive forces to maintain themore » flow rate and velocity gradient at fixed values, while the particles leaving the arteriole at the outlet are removed from the system. Upon valida- tion of this approach, we performed systematic 3D simulations to study plasma skimming in arterioles of diameters 20 to 32 microns. For a flow rate ratio 6:1 at the branches, we observed the \\all-or-nothing" phenomenon with plasma only entering the low flow rate branch. We then simulated blood-plasma separation in arteriolar bifurcations with different bifurcation angles and same diameter of the daughter branches. Our simulations predict a significant increase in RBC flux through the main daughter branch as the bifurcation angle is increased. Lastly, we demonstrated the new methodology for simulating blood flow in ves- sels with multiple inlets and outlets, constructed using an angiogenesis model.« less

Inflow/Outflow Boundary Conditions for Particle-Based Blood Flow Simulations: Application to Arterial Bifurcations and Trees.

PubMed

Lykov, Kirill; Li, Xuejin; Lei, Huan; Pivkin, Igor V; Karniadakis, George Em

2015-08-01

When blood flows through a bifurcation, red blood cells (RBCs) travel into side branches at different hematocrit levels, and it is even possible that all RBCs enter into one branch only, leading to a complete separation of plasma and RBCs. To quantify this phenomenon via particle-based mesoscopic simulations, we developed a general framework for open boundary conditions in multiphase flows that is effective even for high hematocrit levels. The inflow at the inlet is duplicated from a fully developed flow generated in a pilot simulation with periodic boundary conditions. The outflow is controlled by adaptive forces to maintain the flow rate and velocity gradient at fixed values, while the particles leaving the arteriole at the outlet are removed from the system. Upon validation of this approach, we performed systematic 3D simulations to study plasma skimming in arterioles of diameters 20 to 32 microns. For a flow rate ratio 6:1 at the branches, we observed the "all-or-nothing" phenomenon with plasma only entering the low flow rate branch. We then simulated blood-plasma separation in arteriolar bifurcations with different bifurcation angles and same diameter of the daughter branches. Our simulations predict a significant increase in RBC flux through the main daughter branch as the bifurcation angle is increased. Finally, we demonstrated the effectiveness of the new methodology in simulations of blood flow in vessels with multiple inlets and outlets, constructed using an angiogenesis model.

Tabulations of static pressure coefficients on the surfaces of 3 pylon-mounted axisymmetric flow-through nacelles at Mach numbers from 0.40 to 0.98

NASA Technical Reports Server (NTRS)

Re, R. J.; Peddrew, K. H.

1982-01-01

Three flow through nacelles mounted on an 82 deg swept pylon (10 percent thickness-to-chord ratio) were tested in the Langley 16 foot Transonic Tunnel. The long uncambered pylon was supported from a small body of revolution so that pressure measurements on the nacelle and pylon represent a pylon nacelle flow field without a wing present. Two nacelles had NACA 1-85-100 inlets and different circular arc afterbodies. The third nacelle had an NACA 1-70-100 inlet with a circular arc afterbody having the same external shape as one of the other nacelles. Nacelle length to maximum diameter ratio was 3.5. Data were obtained at angles of attack from 2 deg to 8 deg at selected Mach numbers.

Smart actuation of inlet guide vanes for small turbine engine

NASA Astrophysics Data System (ADS)

Rusovici, Razvan; Kwok Choon, Stephen T.; Sepri, Paavo; Feys, Joshuo

2011-04-01

Unmanned Aerial Vehicles (UAVs) have gained popularity over the past few years to become an indispensable part of aerial missions that include reconnaissance, surveillance, and communication [1]. As a result, advancements in small jet-engine performance are needed to increase the performance (range, payload and efficiency) of the UAV. These jet engines designed especially for UAV's are characterized by thrust force on the order of 100N and due to their size and weight limitations, may lack advanced flow control devices such as IGV [2]. The goal of the current study was to present a conceptual design of an IGV smart-material based actuation mechanism that would be simple, compact and lightweight. The compressor section of an engine increases the pressure and conditions the flow before the air enters the combustion chamber [3]. The airflow entering the compressor is often turbulent due to the high angle of incidence between engine inlet and free-stream velocity, or existing atmospheric turbulence. Actuated IGV are used to help control the relative angle of incidence of the flow that enters the engine compressor, thereby preventing flow separation, compressor stall and thus extending the compressor's operating envelope [4]. Turbine jet- engines which employ variable IGV were developed by Rolls Royce (Trent DR-900) and General Electric (J79).

Space shuttle SRM plume expansion sensitivity analysis. [flow characteristics of exhaust gases from solid propellant rocket engines

NASA Technical Reports Server (NTRS)

Smith, S. D.; Tevepaugh, J. A.; Penny, M. M.

1975-01-01

The exhaust plumes of the space shuttle solid rocket motors can have a significant effect on the base pressure and base drag of the shuttle vehicle. A parametric analysis was conducted to assess the sensitivity of the initial plume expansion angle of analytical solid rocket motor flow fields to various analytical input parameters and operating conditions. The results of the analysis are presented and conclusions reached regarding the sensitivity of the initial plume expansion angle to each parameter investigated. Operating conditions parametrically varied were chamber pressure, nozzle inlet angle, nozzle throat radius of curvature ratio and propellant particle loading. Empirical particle parameters investigated were mean size, local drag coefficient and local heat transfer coefficient. Sensitivity of the initial plume expansion angle to gas thermochemistry model and local drag coefficient model assumptions were determined.

Summary of design and blade-element performance data for 12 axial-flow pump rotor configurations

NASA Technical Reports Server (NTRS)

Miller, M. J.; Okiishi, T. H.; Serovy, G. K.; Sandercock, D. M.; Britsch, W. R.

1973-01-01

A collection of noncavitating blade-element performance data for 12 axial-flow pump rotor configurations is presented in tabular form. Rotor design philosophy, test apparatus and procedure, and data reduction and evaluation are discussed. A data storage and recall computer program is described. All but one of the rotor configurations considered were composed of double-circular-arc blade sections and were designed for high inlet relative flow angles. Hub-tip radius ranged from 0.40 to 0.90.

CFD Models of a Serpentine Inlet, Fan, and Nozzle

NASA Technical Reports Server (NTRS)

Chima, R. V.; Arend, D. J.; Castner, R. S.; Slater, J. W.; Truax, P. P.

2010-01-01

Several computational fluid dynamics (CFD) codes were used to analyze the Versatile Integrated Inlet Propulsion Aerodynamics Rig (VIIPAR) located at NASA Glenn Research Center. The rig consists of a serpentine inlet, a rake assembly, inlet guide vanes, a 12-in. diameter tip-turbine driven fan stage, exit rakes or probes, and an exhaust nozzle with a translating centerbody. The analyses were done to develop computational capabilities for modeling inlet/fan interaction and to help interpret experimental data. Three-dimensional Reynolds averaged Navier-Stokes (RANS) calculations of the fan stage were used to predict the operating line of the stage, the effects of leakage from the turbine stream, and the effects of inlet guide vane (IGV) setting angle. Coupled axisymmetric calculations of a bellmouth, fan, and nozzle were used to develop techniques for coupling codes together and to investigate possible effects of the nozzle on the fan. RANS calculations of the serpentine inlet were coupled to Euler calculations of the fan to investigate the complete inlet/fan system. Computed wall static pressures along the inlet centerline agreed reasonably well with experimental data but computed total pressures at the aerodynamic interface plane (AIP) showed significant differences from the data. Inlet distortion was shown to reduce the fan corrected flow and pressure ratio, and was not completely eliminated by passage through the fan

Effect of casing treatment on overall and blade element performance of a compressor rotor

NASA Technical Reports Server (NTRS)

Moore, R. D.; Kovich, G.; Blade, R. J.

1971-01-01

An axial flow compressor rotor was tested at design speed with six different casing treatments across the rotor tip. Radial surveys of pressure, temperature, and flow angle were taken at the rotor inlet and outlet. Surveys were taken at several weight flows for each treatment. All the casings treatments decreased the weight flow at stall over that for the solid casing. Radial surveys indicate that the performance over the entire radial span of the blade is affected by the treatment across the rotor tip.

Two-dimensional compressible flow in centrifugal compressors with straight blades

NASA Technical Reports Server (NTRS)

Stanitz, John D; Ellis, Gaylord O

1950-01-01

Six numerical examples are presented for steady, two-dimensional, compressible, nonviscous flow in centrifugal compressors with thin straight blades, the center lines of which generate the surface of a right circular cone when rotated about the axis of the compressor. A seventh example is presented for incompressible flow. The solutions were obtained in a region of the compressors, including the impeller tip, that was considered to be unaffected by the diffuser vanes or by the impeller-inlet configuration. Each solution applies to radial and mixed flow compressors with various cone angles but with the same angle between blades on the conic flow surface. The solution also apply to radial and mixed flow turbines with the rotation and the flow direction reversed. The effects of variations in the following parameters were investigated: (1) flow rate, (2) impeller-tip speed, (3) variation of passage height with radius, and (4) angle between blades on conic flow surface. The numerical results are presented in plots of the streamlines and constant Mach number lines. Correlation equations are developed whereby the flow conditions in any impeller with straight blades can be determined (in the region investigated by this analysis) for all operating conditions.

Modular jet impingement assemblies with passive and active flow control for electronics cooling

DOEpatents

Zhou, Feng; Dede, Ercan Mehmet; Joshi, Shailesh

2016-09-13

Power electronics modules having modular jet impingement assembly utilized to cool heat generating devices are disclosed. The modular jet impingement assemblies include a modular manifold having a distribution recess, one or more angled inlet connection tubes positioned at an inlet end of the modular manifold that fluidly couple the inlet tube to the distribution recess and one or more outlet connection tubes positioned at an outlet end of the modular manifold that fluidly coupling the outlet tube to the distribution recess. The modular jet impingement assemblies include a manifold insert removably positioned within the distribution recess and include one or more inlet branch channels each including an impinging slot and one or more outlet branch channels each including a collecting slot. Further a heat transfer plate coupled to the modular manifold, the heat transfer plate comprising an impingement surface including an array of fins that extend toward the manifold insert.

A throat-bypass stability system for a YF-12 aircraft research inlet using self-acting mechanical valves

NASA Technical Reports Server (NTRS)

Cole, G. L.; Dustin, M. O.; Neiner, G. H.

1975-01-01

Results of a wind tunnel investigation are presented. The inlet was modified so that airflow can be removed through a porous cowl-bleed region in the vicinity of the throat. Bleed plenum exit flow area is controlled by relief type mechanical valves. Unlike valves in previous systems, these are made for use in a high Mach flight environment and include refinements so that the system could be tested on a NASA YF-12 aircraft. The valves were designed to provide their own reference pressure. The results show that the system can absorb internal-airflow-transients that are too fast for a conventional bypass door control system and that the two systems complement each other quite well. Increased tolerance to angle of attack and Mach number changes is indicated. The valves should provide sufficient time for the inlet control system to make geometry changes required to keep the inlet started.

Numerical investigation of the effects of rising angle on intermediate turbine duct and nearby turbines

NASA Astrophysics Data System (ADS)

Liu, Hongrui; Ji, Lucheng; Liu, Jun; Du, Qiang; Liu, Guang; Wang, Pei; Du, Meimei

2017-10-01

In order to improve the efficiency, ultra-high bypass ratio engine attracts more and more attention because of its huge advantage, which has larger diameter low pressure turbine (LPT). This trend will lead to aggressive (high diffusion) intermediate turbine duct (ITD) design. It is necessary to guide the flow leaving high pressure turbine (HPT) to LPT at a larger diameter without any severe loss generating separation or flow disturbances. In this paper, eight ITDs with upstream swirl vanes and downstream LPT nozzle are investigated with the aid of numerical method. These models are modified from a unique ITD prototype, which comes from a real engine. Key parameters like area ratio, inlet height, and non-dimensional length of the ITDs are kept unchanged, while the rising angle (radial offset) is the only changed parameter which ranges from 8 degrees to 45 degrees. In this paper, the effects of rising angle (RA) on ITD, as well as nearby turbines, will be analyzed in detail. According to the investigation results, RA could be as large as 40 degrees in such model of this paper to escape separation; When RA increases, local inlet flow field of LPT nozzle appears to be with apparent variation; while a positive result is that outlet flow field could be kept almost unchanged through modifying blade profile. On the other hand, it seems optimistic that the overall total pressure loss could be kept nearly equivalent among different RA cases. And a valuable conclusion is that outer wall curvature is more important for pressure loss, which advises a clear direction for optimizing ITD.

An Assessment of the Use of Antimisting Fuel in Turbofan Engines.

DTIC Science & Technology

1981-06-01

Angle PLA Shutoff Lever Angle SOLA Control Speed Nc Compressor Discharge Pressure Ps4 or Pb Compr’ssor Inlet Temperature Tt2 Metered Fuel Flow Wf Control...this comparison the Royal Aircraft Establishment deqrader had a lower filter ratio reduction, consumed more power, and had a higher tempera- ture rise...negligible. This would imply that little of the total enerqy consume 1 by the pump goes towa~rds d(,qrading the antimisting kerosene. Further dita analysis

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

NASA Technical Reports Server (NTRS)

Head, V. L.

1972-01-01

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Effect of Boundary-Layer Bleed Hole Inclination Angle and Scaling on Flow Coefficient Behavior

NASA Technical Reports Server (NTRS)

Eichorn, Michael B.; Barnhart, Paul J.; Davis, David O.; Vyas, Manan A.; Slater, John W.

2013-01-01

Phase II data results of the Fundamental Inlet Bleed Experiments study at NASA Glenn Research Center are presented which include flow coefficient behavior for 21 bleed hole configurations. The bleed configurations are all round holes with hole diameters ranging from 0.795 to 6.35 mm, hole inclination angles from 20deg to 90deg, and thickness-to-diameter ratios from 0.25 to 2.0. All configurations were tested at a unit Reynolds number of 2.46 10(exp 7)/m and at discrete local Mach numbers of 1.33, 1.62, 1.98, 2.46, and 2.92. Interactions between the design parameters of hole diameter, hole inclination angle, and thickness-to-diameter as well as the interactions between the flow parameters of pressure ratio and Mach number upon the flow coefficient are examined, and a preliminary statistical model is proposed. An existing correlation is also examined with respect to this data.

Spray Characteristics of a Hybrid Twin-Fluid Pressure-Swirl Atomizer

NASA Technical Reports Server (NTRS)

Durham, M. J.; Sojka, P. E.; Ashmore, C. B.

2004-01-01

The spray performance of a fuel injection system applicable for use in main combustion chamber of an oxidizer-rich staged combustion (ORSC) cycles is presented. The experimental data reported here include mean drop size and drop size distribution, spray cone half-angle, and momentum rate (directly related to spray penetration). The maximum entropy formalism, MEF, method to predict drop size distribution is applied and compared to the experimental data. Geometric variables considered include the radius of the injector inlet orifice plate through which oxidizer flows (&) and the exposed length from the fuel inlet to the injector exit plane (L2). Operating conditions that were varied include the liquid mass flow rate and air mass flow rate. For orifices B and C there is a significant dependence of D3Z on both the air and liquid mass flow rates, as well as on L2. For the A orifice, the momentum rate of the air flow appears to exceed a threshold value above which a constant D32 is obtained. Using the MEF method, a semi-analytical process was developed to model the spray distribution using two input parameters (q = 0.4 and Dso). The momentum rate of the spray is directly related to the air and liquid mass flow rates. The cone half angle of the spray ranges from 25 to 17 degrees. The data resulting from this project will eventually be used to develop advanced rocket systems.

Flight Test Results from the Rake Airflow Gage Experiment on the F-15B Airplane

NASA Technical Reports Server (NTRS)

Frederick, Michael A.; Ratnayake, Nalin A.

2011-01-01

The Rake Airflow Gage Experiment involves a flow-field survey rake that was flown on the Propulsion Flight Test Fixture at the NASA Dryden Flight Research Center using the Dryden F-15B research test bed airplane. The objective of this flight test was to ascertain the flow-field angularity, local Mach number profile, total pressure distortion, and dynamic pressure at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment. This new mixed-compression, supersonic inlet is planned for flight test in the near term. Knowledge of the flow-field characteristics at this location underneath the airplane is essential to flight test planning and computational modeling of the new inlet, an< it is also applicable for future propulsion systems research that may use the Propulsion Flight Test Fixture. This report describes the flight test preparation and execution, and the local flow-field properties calculated from pressure measurements of the rake. Data from the two Rake Airflow Gage Experiment research flights demonstrate that the F-15B airplane, flying at a free-stream Mach number of 1.65 and a pressure altitude of 40,000 ft, would achieve the desired local Mach number for the future inlet flight test. Interface plane distortion levels of 2 percent and a local angle of attack of -2 deg were observed at this condition. Alternative flight conditions for future testing and an exploration of certain anomalous data also are provided.

The prediction of noise and installation effects of high-subsonic dual-stream jets in flight

NASA Astrophysics Data System (ADS)

Saxena, Swati

Both military and civil aircraft in service generate high levels of noise. One of the major contributors to this noise generated from the aircraft is the jet engine exhaust. This makes the study of jet noise and methods to reduce jet noise an active research area with the aim of designing quieter military and commercial aircraft. The current stringent aircraft noise regulations imposed by the Federal Aviation Administration (FAA) and other international agencies, have further raised the need to perform accurate jet noise calculations for more reliable estimation of the jet noise sources. The main aim of the present research is to perform jet noise simulations of single and dual-stream jets with engineering accuracy and assess forward flight effects on the jet noise. Installation effects such as caused by the pylon are also studied using a simplified pylon nozzle configuration. Due to advances in computational power, it has become possible to perform turbulent flow simulations of high speed jets, which leads to more accurate noise predictions. In the present research, a hybrid unsteady RANS-LES parallel multi-block structured grid solver called EAGLEJet is written to perform the nozzle flow calculations. The far-field noise calculation is performed using solutions to the Ffowcs Williams and Hawkings equation. The present calculations use meshes with 5 to 11 million grid points and require about three weeks of computing time with about 100 processors. A baseline single stream convergent nozzle and a dual-stream coaxial convergent nozzle are used for the flow and noise analysis. Calculations for the convergent nozzle are performed at a high subsonic jet Mach number of Mj = 0.9, which is similar to the operating conditions for commercial aircraft engines. A parallel flow gives the flight effect, which is simulated with a co-flow Mach number, Mcf varying from 0.0 to 0.28. The grid resolution effects, statistical properties of the turbulence and the heated jet effects ( TTR = 2.7) are studied and related to the noise characteristics of the jet. Both flow and noise predictions show good agreement with PIV and microphone measurements. The potential core lengths and nozzle wall boundary characteristics are studied to understand the differences between the numerical potential core lengths as compared to experiments. The flight velocity exponent, m is calculated from the noise reduction in overall sound pressure levels (OASPL, dB) and relative velocity (V j -- Vcf) at all jet inlet (angular) angles. The variation of the exponent, m at lower (50° to 90°) and higher aft inlet angles (120° to 150°) is studied and compared with available measurements. Previous studies have shown a different variation of the exponent with inlet angles while the current numerical data match well with recent experiments conducted on the same nozzle geometry. Today, turbofans are the most efficient engines in service used in almost all major commercial aircraft. Turbofans have a dual-stream exhaust nozzle with primary and secondary flow whose flow and noise characteristics are different from that of single stream jets. A Boeing-designed coaxial nozzle, with area ratio of As/Ap = 3.0, is used to study dual-stream jet noise in the present research. In this configuration, the primary nozzle extends beyond the secondary nozzle, which is representative of large turbofan engines in commercial service. The flow calculations are performed at high subsonic Mach numbers in the primary and secondary nozzles (Mpj = 0.85, Msj = 0.95) with heated core flow, TTRp = 2.26 and unheated fan flow, TTRs = 1.0. The co-flow of Mcf = 0.2 is used. The subscript p, s and amb represent the primary (core) nozzle, the secondary (fan) nozzle, and the ambient flow conditions, respectively. The statistical properties in the primary and secondary shear layers are studied and compared with those of the single stream jets. It has been found that the eddy convection velocity is lower in dual-stream jets as compared to the single stream jet operating at a similar jet exit Mach number. The phase velocity is higher in the secondary shear layer as compared to primary shear layer. The noise measurements agree well with the predicted data and noise reduction is observed in the presence of co-flow. The variation of the flight velocity exponent is calculated as a function of nozzle inlet angle. The value of the exponent at higher inlet angles is lower as compared to the single stream jets. This suggests that the noise levels are less affected in the peak noise direction in the presence of co-flow in dual-stream jets as compared to single stream jets. Two reference velocities: primary jet exit velocity Vpj and mixed velocity Vmix are considered which result in different absolute values of the exponents. Scaling of the jet spectra is performed at different inlet angles and good collapse has been obtained between the spectra. The installation effects on jet noise are studied using a simplified pylon structure with a dual-stream nozzle. In the presence of a pylon, the azimuthal symmetry of the nozzle is lost and thus the flow characteristics are different as compared to the baseline nozzle. This will result in different noise characteristics of the installed jet.

A throat-bypass stability-bleed system using relief valves to increase the transient stability of a mixed-compression inlet. [YF-12 aircraft inlet tests in the Lewis 10 by 10 ft supersonic wind tunnel

NASA Technical Reports Server (NTRS)

Neiner, G. H.; Dustin, M. O.; Cole, G. L.

1979-01-01

A stability-bleed system was installed in a YF-12 flight inlet that was subjected to internal and external airflow disturbances in the NASA Lewis 10 by 10 foot supersonic wind tunnel. The purpose of the system is to allow higher inlet performance while maintaining a substantial tolerance (without unstart) to internal and external disturbances. At Mach numbers of 2.47 and 2.76, the inlet tolerance to decreases in diffuser-exit corrected airflow was increased by approximately 10 percent of the operating-point airflow. The stability-bleed system complemented the terminal-shock-control system of the inlet and did not show interaction problems. For disturbances which caused a combined decrease in Mach number and increase in angle of attack, the system with valves operative kept the inlet started 4 to 28 times longer than with the valves inoperative. Hence, the stability system provides additional time for the inlet control system to react and prevent unstart. This was observed for initial Mach numbers of 2.55 and 2.68. For slow increase in angle of attack at Mach 2.47 and 2.76, the system kept the inlet started beyond the steady-state unstart angle. However, the maximum transient angles of attack without unstart could not be determined because wind-tunnel mechanical-stop limits for angle of attack were reached.

Hypersonic Inlet for a Laser Powered Propulsion System

NASA Astrophysics Data System (ADS)

Harrland, Alan; Doolan, Con; Wheatley, Vincent; Froning, Dave

2011-11-01

Propulsion within the lightcraft concept is produced via laser induced detonation of an incoming hypersonic air stream. This process requires suitable engine configurations that offer good performance over all flight speeds and angles of attack to ensure the required thrust is maintained. Stream traced hypersonic inlets have demonstrated the required performance in conventional hydrocarbon fuelled scramjet engines, and has been applied to the laser powered lightcraft vehicle. This paper will outline the current methodology employed in the inlet design, with a particular focus on the performance of the lightcraft inlet at angles of attack. Fully three-dimensional turbulent computational fluid dynamics simulations have been performed on a variety of inlet configurations. The performance of the lightcraft inlets have been evaluated at differing angles of attack. An idealized laser detonation simulation has also been performed to validate that the lightcraft inlet does not unstart during the laser powered propulsion cycle.

FLOW FIELD IN SUPERSONIC MIXED-COMPRESSION INLETS AT ANGLE OF ATTACK USING THE THREE DIMENSIONAL METHOD OF CHARACTERISTICS WITH DISCRETE SHOCK WAVE FITTING

NASA Technical Reports Server (NTRS)

Bishop, A. R.

1994-01-01

This computer program calculates the flow field in the supersonic portion of a mixed-compression aircraft inlet at non-zero angle of attack. This approach is based on the method of characteristics for steady three-dimensional flow. The results of this program agree with those produced by the two-dimensional method of characteristics when axisymmetric flow fields are calculated. Except in regions of high viscous interaction and boundary layer removal, the results agree well with experimental data obtained for threedimensional flow fields. The flow field in a variety of axisymmetric mixed compression inlets can be calculated using this program. The bow shock wave and the internal shock wave system are calculated using a discrete shock wave fitting procedure. The internal flow field can be calculated either with or without the discrete fitting of the internal shock wave system. The influence of molecular transport can be included in the calculation of the external flow about the forebody and in the calculation of the internal flow when internal shock waves are not discretely fitted. The viscous and thermal diffussion effects are included by treating them as correction terms in the method of characteristics procedure. Dynamic viscosity is represented by Sutherland's law and thermal conductivity is represented as a quadratic function of temperature. The thermodynamic model used is that of a thermally and calorically perfect gas. The program assumes that the cowl lip is contained in a constant plane and that the centerbody contour and cowl contour are smooth and have continuous first partial derivatives. This program cannot calculate subsonic flow, the external flow field if the bow shock wave does not exist entirely around the forebody, or the internal flow field if the bow flow field is injected into the annulus. Input to the program consists of parameters to control execution, to define the geometry, and the vehicle orientation. Output consists of a list of parameters used, solution planes, and a description of the shock waves. This program is written in FORTRAN IV for batch execution and has been implemented on a CDC 6000 series machine with a central memory requirement of 110K (octal) of 60 bit words when it is overlayed. This flow analysis program was developed in 1978.

Heat Transfer Measurements and Predictions on a Power Generation Gas Turbine Blade

NASA Technical Reports Server (NTRS)

Giel, Paul W.; Bunker, Ronald S.; VanFossen, G. James; Boyle, Robert J.

2000-01-01

Detailed heat transfer measurements and predictions are given for a power generation turbine rotor with 129 deg of nominal turning and an axial chord of 137 mm. Data were obtained for a set of four exit Reynolds numbers comprised of the design point of 628,000, -20%, +20%, and +40%. Three ideal exit pressure ratios were examined including the design point of 1.378, -10%, and +10%. Inlet incidence angles of 0 deg and +/-2 deg were also examined. Measurements were made in a linear cascade with highly three-dimensional blade passage flows that resulted from the high flow turning and thick inlet boundary layers. Inlet turbulence was generated with a blown square bar grid. The purpose of the work is the extension of three-dimensional predictive modeling capability for airfoil external heat transfer to engine specific conditions including blade shape, Reynolds numbers, and Mach numbers. Data were obtained by a steady-state technique using a thin-foil heater wrapped around a low thermal conductivity blade. Surface temperatures were measured using calibrated liquid crystals. The results show the effects of strong secondary vortical flows, laminar-to-turbulent transition, and also show good detail in the stagnation region.

Design-of-experiments to Reduce Life-cycle Costs in Combat Aircraft Inlets

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Baust, Henry D.; Agrell, Johan

2003-01-01

It is the purpose of this study to demonstrate the viability and economy of Design- of-Experiments (DOE), to arrive at micro-secondary flow control installation designs that achieve optimal inlet performance for different mission strategies. These statistical design concepts were used to investigate the properties of "low unit strength" micro-effector installation. "Low unit strength" micro-effectors are micro-vanes, set a very low angle-of incidence, with very long chord lengths. They are designed to influence the neat wall inlet flow over an extended streamwise distance. In this study, however, the long chord lengths were replicated by a series of short chord length effectors arranged in series over multiple bands of effectors. In order to properly evaluate the performance differences between the single band extended chord length installation designs and the segmented multiband short chord length designs, both sets of installations must be optimal. Critical to achieving optimal micro-secondary flow control installation designs is the understanding of the factor interactions that occur between the multiple bands of micro-scale vane effectors. These factor interactions are best understood and brought together in an optimal manner through a structured DOE process, or more specifically Response Surface Methods (RSM).

V/STOL Tandem Fan transition section model test. [in the Lewis Research Center 10-by-10 foot wind tunnel

NASA Technical Reports Server (NTRS)

Simpkin, W. E.

1982-01-01

An approximately 0.25 scale model of the transition section of a tandem fan variable cycle engine nacelle was tested in the NASA Lewis Research Center 10-by-10 foot wind tunnel. Two 12-inch, tip-turbine driven fans were used to simulate a tandem fan engine. Three testing modes simulated a V/STOL tandem fan airplane. Parallel mode has two separate propulsion streams for maximum low speed performance. A front inlet, fan, and downward vectorable nozzle forms one stream. An auxilliary top inlet provides air to the aft fan - supplying the core engine and aft vectorable nozzle. Front nozzle and top inlet closure, and removal of a blocker door separating the two streams configures the tandem fan for series mode operations as a typical aircraft propulsion system. Transition mode operation is formed by intermediate settings of the front nozzle, blocker door, and top inlet. Emphasis was on the total pressure recovery and flow distortion at the aft fan face. A range of fan flow rates were tested at tunnel airspeeds from 0 to 240 knots, and angles-of-attack from -10 to 40 deg for all three modes. In addition to the model variables for the three modes, model variants of the top inlet were tested in the parallel mode only. These lip variables were: aft lip boundary layer bleed holes, and Three position turning vane. Also a bellmouth extension of the top inlet side lips was tested in parallel mode.

Computer program for design analysis of radial-inflow turbines

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1976-01-01

A computer program written in FORTRAN that may be used for the design analysis of radial-inflow turbines was documented. The following information is included: loss model (estimation of losses), the analysis equations, a description of the input and output data, the FORTRAN program listing and list of variables, and sample cases. The input design requirements include the power, mass flow rate, inlet temperature and pressure, and rotational speed. The program output data includes various diameters, efficiencies, temperatures, pressures, velocities, and flow angles for the appropriate calculation stations. The design variables include the stator-exit angle, rotor radius ratios, and rotor-exit tangential velocity distribution. The losses are determined by an internal loss model.

Analytical study on a two-dimensional plane of the off-design flow properties of tandem-bladed compressor stators

NASA Technical Reports Server (NTRS)

Sanger, N. L.

1973-01-01

The flow characteristics of several tandem bladed compressor stators were analytically evaluated over a range of inlet incidence angles. The ratios of rear-segment to front-segment chord and camber were varied. Results were also compared to the analytical performance of a reference solid blade section. All tandem blade sections exhibited lower calculated losses than the solid stator. But no one geometric configuration exhibited clearly superior characteristics. The front segment accepts the major effect of overall incidence angle change. Rear- to front-segment camber ratios of 4 and greater appeared to be limited by boundary-layer separation from the pressure surface of the rear segment.

Numerical investigation on pressure fluctuations in centrifugal compressor with different inlet guide vanes pre-whirl angles

NASA Astrophysics Data System (ADS)

Wang, Y. C.; Shi, M.; Cao, S. L.; Li, Z. H.

2013-12-01

The pressure fluctuations in a centrifugal compressor with different inlet guide vanes (IGV) pre-whirl angles were investigated numerically, as well as the pre-stress model and static structural of blade. The natural frequency was evaluated by pre-stress model analysis. The results show that, the aero-dynamic pressure acting on blade surface is smaller than rotation pre-stress, which wouldn't result in large deformation of blade. The natural frequencies with rotation pre-stress are slightly higher than without rotation pre-stress. The leading mechanism of pressure fluctuations for normal conditions is the rotor-stator (IGVs) interaction, while is serious flow separations for conditions that are close to surge line. A few frequency components in spectra are close to natural frequency, which possibly result in resonant vibration if amplitude is large enough, which is dangerous for compressor working, and should be avoided.

Algorithm for Controlling a Centrifugal Compressor

NASA Technical Reports Server (NTRS)

Benedict, Scott M.

2004-01-01

An algorithm has been developed for controlling a centrifugal compressor that serves as the prime mover in a heatpump system. Experimental studies have shown that the operating conditions for maximum compressor efficiency are close to the boundary beyond which surge occurs. Compressor surge is a destructive condition in which there are instantaneous reversals of flow associated with a high outlet-to-inlet pressure differential. For a given cooling load, the algorithm sets the compressor speed at the lowest possible value while adjusting the inlet guide vane angle and diffuser vane angle to maximize efficiency, subject to an overriding requirement to prevent surge. The onset of surge is detected via the onset of oscillations of the electric current supplied to the compressor motor, associated with surge-induced oscillations of the torque exerted by and on the compressor rotor. The algorithm can be implemented in any of several computer languages.

Internal combustion engine system having a power turbine with a broad efficiency range

DOEpatents

Whiting, Todd Mathew; Vuk, Carl Thomas

2010-04-13

An engine system incorporating an air breathing, reciprocating internal combustion engine having an inlet for air and an exhaust for products of combustion. A centripetal turbine receives products of the combustion and has a housing in which a turbine wheel is rotatable. The housing has first and second passages leading from the inlet to discrete, approximately 180.degree., portions of the circumference of the turbine wheel. The passages have fixed vanes adjacent the periphery of the turbine wheel and the angle of the vanes in one of the passages is different than those in the other so as to accommodate different power levels providing optimum approach angles between the gases passing the vanes and the blades of the turbine wheel. Flow through the passages is controlled by a flapper valve to direct it to one or the other or both passages depending upon the load factor for the engine.

Low speed and angle of attack effects on sonic and near-sonic inlets

NASA Technical Reports Server (NTRS)

Hickcox, T. E.; Lawrence, R. L.; Syberg, J.; Wiley, D. R.

1975-01-01

Tests of the Quiet, Clean Short-Haul Experimental Engine (QCSEE) were conducted to determine the effects of forward velocity and angle of attack on sonic and near-sonic inlet aerodynamic performance penalties and acoustic suppression characteristics. The tests demonstrate that translating centerbody and radial vane sonic inlets, and QCSEE high throat Mach number inlets, can be designed to operate effectively at forward speed and moderate angle of attack with good performance and noise suppression capability. The test equipment and procedures used in conducting the evaluation are described. Results of the tests are presented in tabular form.

Micro-Ramps for External Compression Low-Boom Inlets

NASA Technical Reports Server (NTRS)

Rybalko, Michael; Loth, Eric; Chima, Rodrick V.; Hirt, Stefanie M.; DeBonis, James R.

2010-01-01

The application of vortex generators for flow control in an external compression, axisymmetric, low-boom concept inlet was investigated using RANS simulations with three-dimensional (3-D), structured, chimera (overset) grids and the WIND-US code. The low-boom inlet design is based on previous scale model 1- by 1-ft wind tunnel tests and features a zero-angle cowl and relaxed isentropic compression centerbody spike, resulting in defocused oblique shocks and a weak terminating normal shock. Validation of the methodology was first performed for micro-ramps in supersonic flow on a flat plate with and without oblique shocks. For the inlet configuration, simulations with several types of vortex generators were conducted for positions both upstream and downstream of the terminating normal shock. The performance parameters included incompressible axisymmetric shape factor, separation area, inlet pressure recovery, and massflow 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. The optimum upstream configuration was found to substantially reduce the post-shock separation area but did not significantly impact recovery at the aerodynamic interface plane (AIP). Downstream device placement allowed for fuller boundary layer velocity profiles and reduced distortion. This resulted in an improved pressure recovery and massflow ratio at the AIP compared to the baseline solid-wall configuration.

Inlet flowfield investigation. Part 2: Computation of the flow about a supercruise forebody at supersonic speeds

NASA Technical Reports Server (NTRS)

Paynter, G. C.; Salemann, V.; Strom, E. E. I.

1984-01-01

A numerical procedure which solves the parabolized Navier-Stokes (PNS) equations on a body fitted mesh was used to compute the flow about the forebody of an advanced tactical supercruise fighter configuration in an effort to explore the use of a PNS method for design of supersonic cruise forebody geometries. Forebody flow fields were computed at Mach numbers of 1.5, 2.0, and 2.5, and at angles-of-attack of 0 deg, 4 deg, and 8 deg. at each Mach number. Computed results are presented at several body stations and include contour plots of Mach number, total pressure, upwash angle, sidewash angle and cross-plane velocity. The computational analysis procedure was found reliable for evaluating forebody flow fields of advanced aircraft configurations for flight conditions where the vortex shed from the wing leading edge is not a dominant flow phenomenon. Static pressure distributions and boundary layer profiles on the forebody and wing were surveyed in a wind tunnel test, and the analytical results are compared to the data. The current status of the parabolized flow flow field code is described along with desirable improvements in the code.

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.)

Deflected jet experiments in a turbulent combustor flowfield. Ph.D. Thesis Final Report

NASA Technical Reports Server (NTRS)

Ferrell, G. B.; Lilley, D. G.

1985-01-01

Experiments were conducted to characterize the time-mean and turbulent flow field of a deflected turbulent jet in a confining cylindrical crossflow. Jet-to-crossflow velocity ratios of 2, 4, and 6 were investigated, under crossflow inlet swirler vane angles of 0 (swirler removed), 45 and 70 degrees. Smoke, neutrally buoyant helium-filled soap bubbles, and multi-spark flow visualization were employed to highlight interesting features of the deflected jet, as well as the tracjectory and spread pattern of the jet. A six-position single hot-wire technique was used to measure the velocities and turbulent stresses in nonswirling crossflow cases. In these cases, measurements confirmed that the deflected jet is symmetrical about the vertical plan passing through the crossflow axis, and the jet penetration was found to be reduced from that of comparable velocity ratio infinite crossflow cases. In the swirling crossflow cases, the flow visualization techniques enabled gross flow field characterization to be obtained for a range of lateral jet-to-crossflow velocity ratios and a range of inlet swirl strengths in the main flow.

Numerical study of centrifugal compressor stage vaneless diffusers

NASA Astrophysics Data System (ADS)

Galerkin, Y.; Soldatova, K.; Solovieva, O.

2015-08-01

The authors analyzed CFD calculations of flow in vaneless diffusers with relative width in range from 0.014 to 0.100 at inlet flow angles in range from 100 to 450 with different inlet velocity coefficients, Reynolds numbers and surface roughness. The aim is to simulate calculated performances by simple algebraic equations. The friction coefficient that represents head losses as friction losses is proposed for simulation. The friction coefficient and loss coefficient are directly connected by simple equation. The advantage is that friction coefficient changes comparatively little in range of studied parameters. Simple equations for this coefficient are proposed by the authors. The simulation accuracy is sufficient for practical calculations. To create the complete algebraic model of the vaneless diffuser the authors plan to widen this method of modeling to diffusers with different relative length and for wider range of Reynolds numbers.

Experimental evaluation of shockless supercritical airfoils in cascade

NASA Technical Reports Server (NTRS)

Boldman, D. R.; Buggele, A. E.; Shaw, L. M.

1983-01-01

Surface Mach number distributions, total pressure loss coefficients, and schlieren images of the flow are presented over a range of inlet Mach numbers and air angles. Several different trailing edge geometries were tested. At design conditions a leading edge separation bubble was observed resulting in higher losses than anticipated. The minimum losses were obtained at a negative incidence condition in which the flow was accelerating over most of the supercritical region. Relatively minor differences in losses were measured with the different trailing edge geometries studied.

Experimental Verification of the Streamline Curvature Numerical Analysis Method Applied to the Flow through an Axial Flow Fan.

DTIC Science & Technology

1980-05-28

Total Deviation Angles and Measured Inlet Axial Velocity . . . . 55 ix LIST OF FIGURES (Continued) Figure Page 19 Points Defining Blade Sections of...distance from leading edge to point of maximum camber along chord line ar tip vortex core radius AVR axial velocity ratio (Vx /V x c chord length CL tip...yaw ceoefficient d longitudinal distance from leading edge to tip vortex calculation point G distance from chord line to maximum camber point K cascade

Experimental and Theoretical Study of Flow Fields Around Ducted-Nacelle Models

NASA Technical Reports Server (NTRS)

Mack, Robert J.

1998-01-01

The flow field near four small-scale ducted-nacelle bodies of revolution has been analytically and experimentally studied to determine exterior and interior mass-flow characteristics, and to measure flow-field overpressures generated by the nacelle's forebody shape. Four nacelle models with the same profile, but of different sizes, were used in the study. Shadowgraph pictures showed inlet shocks attached to the cowl lip (indicating unchoked flow) on all four models, at all the test Mach numbers, through an angle of attack range of 0.0 to 6.0 degrees. Pressure signatures measured in the flow field of the largest of the four nacelle models were compared with those predicted by corrected and uncorrected Whitham theory. At separation distances greater than 3.0 to 4.0 inlet diameters, good agreement was found. Poorer agreement was found at extreme near-field separation distances, but this was attributed to pressure-gage limitations and probe-flow field interactions. The overall favorable results supported a conclusion that corrected Whitham theory was sufficiently accurate to make the nacelle-wing interference-lift code useful for sonic-boom analysis and the preliminary design of supersonic-cruise conceptual aircraft.

Effects of Swirler Shape on Two-Phase Swirling Flow in a Steam Separator

NASA Astrophysics Data System (ADS)

Kataoka, Hironobu; Shinkai, Yusuke; Tomiyama, Akio

Experiments on two-phase swirling flow in a separator are carried out using several swirlers having different vane angles, different hub diameters and different number of vanes to seek a way for improving steam separators of uprated boiling water reactors. Ratios of the separated liquid flow rate to the total liquid flow rate, flow patterns, liquid film thicknesses and pressure drops are measured to examine the effects of swirler shape on air-water two-phase swirling annular flows in a one-fifth scale model of the separator. As a result, the following conclusions are obtained for the tested swirlers: (1) swirler shape scarcely affects the pressure drop in the barrel of the separator, (2) decreasing the vane angle is an effective way for reducing the pressure drop in the diffuser of the separator, and (3) the film thickness at the inlet of the pick-off-ring of the separator is not sensitive to swirler shape, which explains the reason why the separator performance does not depend on swirler shape.

Super-Cavitating Flow Around Two-Dimensional Conical, Spherical, Disc and Stepped Disc Cavitators

NASA Astrophysics Data System (ADS)

Sooraj, S.; Chandrasekharan, Vaishakh; Robson, Rony S.; Bhanu Prakash, S.

2017-08-01

A super-cavitating object is a high speed submerged object that is designed to initiate a cavitation bubble at the nose which extends past the aft end of the object, substantially reducing the skin friction drag that would be present if the sides of the object were in contact with the liquid in which the object is submerged. By reducing the drag force the thermal energy consumption to move faster can also be minimised. The super-cavitation behavioural changes with respect to Cavitators of various geometries have been studied by varying the inlet velocity. Two-dimensional computational fluid dynamics analysis has been carried out by applying k-ε turbulence model. The variation of drag coefficient, cavity length with respect to cavitation number and inlet velocity are analyzed. Results showed conical Cavitator with wedge angle of 30° has lesser drag coefficient and cavity length when compared to conical Cavitators with wedge angles 45° and 60°, spherical, disc and stepped disc Cavitators. Conical cavitator 60° and disc cavitator have the maximum cavity length but with higher drag coefficient. Also there is significant variation of supercavitation effect observed between inlet velocities of 32 m/s to 40 m/s.

Prediction of active control of subsonic centrifugal compressor rotating stall

NASA Technical Reports Server (NTRS)

Lawless, Patrick B.; Fleeter, Sanford

1993-01-01

A mathematical model is developed to predict the suppression of rotating stall in a centrifugal compressor with a vaned diffuser. This model is based on the employment of a control vortical waveform generated upstream of the impeller inlet to damp weak potential disturbances that are the early stages of rotating stall. The control system 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. The model was effective at predicting the stalling behavior of the Purdue Low Speed Centrifugal Compressor for two distinctly different stall patterns. Predictions made for the effect of a controlled inlet vorticity wave on the stability of the compressor show that for minimum control wave magnitudes, on the order of the total inlet disturbance magnitude, significant damping of the instability can be achieved. For control waves of sufficient amplitude, the control phase angle appears to be the most important factor in maintaining a stable condition in the compressor.

Influence of nanofluids on the efficiency of Flat-Plate Solar Collectors (FPSC)

NASA Astrophysics Data System (ADS)

Nejad, Marjan B.; Mohammed, H. A.; Sadeghi, O.; Zubeer, Swar A.

2017-11-01

A numerical investigation is performed using finite volume method to study the laminar heat transfer in a three-dimensional flat-plate solar collector using different nanofluids as working fluids. Three nanofluids with different types of nanoparticles (Ag, MWCNT and Al2O3 dispersed in water) with 1-2 wt% volume fractions are analyzed. A constant heat flux, equivalent to solar radiation absorbed by the collector, is applied at the top surface of the absorber plate. In this study, several parameters including boundary conditions (different volume flow rates, different fluid inlet temperatures and different solar irradiance at Skudai, Malaysia), different types of nanoparticles, and different solar collector tilt angles are investigated to identify their effects on the heat transfer performance of FPSC. The numerical results reveal that the three types of nanofluid enhance the thermal performance of solar collector compared to pure water and FPSC with Ag nanofluid has the best thermal performance enhancement. For all the cases, the collector efficiency increased with the increase of volume flow rate while fluid outlet temperature decreased. It is found that FPSC with tilt angle of 10° and fluid inlet temperature of 301.15 K has the best thermal performance.

Kinetics of gravity-driven slug flow in partially wettable capillaries of varying cross section

NASA Astrophysics Data System (ADS)

Nissan, Alon; Wang, Qiuling; Wallach, Rony

2016-11-01

A mathematical model for slug (finite liquid volume) motion in not-fully-wettable capillary tubes with sinusoidally varying cross-sectional areas was developed. The model, based on the Navier-Stokes equation, accounts for the full viscous terms due to nonuniform geometry, the inertial term, the slug's front and rear meniscus hysteresis effect, and dependence of contact angle on flow velocity (dynamic contact angle). The model includes a velocity-dependent film that is left behind the advancing slug, reducing its mass. The model was successfully verified experimentally by recording slug movement in uniform and sinusoidal capillary tubes with a gray-scale high-speed camera. Simulation showed that tube nonuniformity has a substantial effect on slug flow pattern: in a uniform tube it is monotonic and depends mainly on the slug's momentary mass/length; an undulating tube radius results in nonmonotonic flow characteristics. The static nonzero contact angle varies locally in nonuniform tubes owing to the additional effect of wall slope. Moreover, the nonuniform cross-sectional area induces slug acceleration, deceleration, blockage, and metastable-equilibrium locations. Increasing contact angle further amplifies the geometry effect on slug propagation. The developed model provides a modified means of emulating slug flow in differently wettable porous media for intermittent inlet water supply (e.g., raindrops on the soil surface).

Effects of radial diffuser hydraulic design on a double-suction centrifugal pump

NASA Astrophysics Data System (ADS)

Hou, H. C.; Zhang, Y. X.; Xu, C.; Zhang, J. Y.; Li, Z. L.

2016-05-01

In order to study effects of radial diffuser on hydraulic performance of crude oil pump, the steady CFD numerical method is applied and one large double-suction oil pump running in long-distance pipeline is considered. The research focuses on analysing the influence of its diffuser vane profile on hydraulic performance of oil pump. The four different types of cylindrical vane have been designed by in-house codes mainly including double arcs (DA), triple arcs (TA), equiangular spiral line (ES) and linear variable angle spiral line (LVS). During design process diffuser vane angles at inlet and outlet are tentatively given within a certain range and then the wrapping angle of the four types of diffuser vanes can be calculated automatically. Under the given inlet and outlet angles, the linear variable angle spiral line profile has the biggest wrapping angle and profile length which is good to delay channel diffusion but bring more friction hydraulic loss. Finally the vane camber line is thickened at the certain uniform thickness distribution and the 3D diffuser models are generated. The whole flow passage of oil pump with different types of diffusers under various flow rate conditions are numerically simulated based on RNG k-ɛ turbulent model and SIMPLEC algorithm. The numerical results show that different types of diffusers can bring about great difference on the hydraulic performance of oil pump, of which the ES profile diffuser with its proper setting angle shows the best hydraulic performance and its inner flow field is improved obviously. Compared with the head data from model sample, all designed diffusers can make a certain improvement on head characteristic. At the large flow rate conditions the hydraulic efficiency increases obviously and the best efficiency point shift to the large flow rate range. The ES profile diffuser embodies the better advantages on pump performance which can be explained theoretically that the diffuser actually acts as a diffusion device and is good to transform the dynamic energy to pressure energy. Then through the hydraulic loss analysis of each pump component for all diffusers, it shows that the impeller takes up the biggest part of the whole loss about 8.19% averagely, the radial diffuser about 3.70% and the volute about 1.65%. The hydraulic loss of impeller is dominant at the large flow rate while the radial diffuser is at the small flow rate. Among all diffusers, the ES profile diffuser generates the least loss and combined to the distribution of velocity vector and turbulent kinetic energy for two kinds of diffusers it also shows that ES profile is fit to apply in radial diffuser. This research can offer a significant reference for the radial diffuser hydraulic design of such centrifugal pumps.

Flight Test Results from the Rake Airflow Gage Experiment on the F-15B Airplane

NASA Technical Reports Server (NTRS)

Frederick, Michael A.; Ratnayake, Nalin A.

2010-01-01

The Rake Airflow Gage Experiment involves a flow-field survey rake that was flown on the Propulsion Flight Test Fixture at the NASA Dryden Flight Research Center using the Dryden F-15B research test bed airplane. The objective of this flight test was to ascertain the flow-field angularity, local Mach number profile, total pressure distortion, and dynamic pressure at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment. This new mixed-compression, supersonic inlet is planned for flight test in the near term. Knowledge of the flow-field characteristics at this location underneath the airplane is essential to flight test planning and computational modeling of the new inlet, and it is also applicable for future propulsion systems research that may use the Propulsion Flight Test Fixture. This report describes the flight test preparation and execution, and the local flowfield properties calculated from pressure measurements of the rake. Data from the two Rake Airflow Gage Experiment research flights demonstrate that the F-15B airplane, flying at a free-stream Mach number of 1.65 and a pressure altitude of 40,000 ft, would achieve the desired local Mach number for the future inlet flight test. Interface plane distortion levels of 2 percent and a local angle of attack of 2 were observed at this condition. Alternative flight conditions for future testing and an exploration of certain anomalous data also are provided.

An Experimental Device for Generating High Frequency Perturbations in Supersonic Wind Tunnels

NASA Technical Reports Server (NTRS)

Melcher, Kevin J.; Ibrahim, Mounir B.

1996-01-01

This paper describes the analytical study of a device that has been proposed as a mechanism for generating gust-like perturbations in supersonic wind tunnels. The device is envisioned as a means to experimentally validate dynamic models and control systems designed for high-speed inlets. The proposed gust generator is composed of two flat trapezoidal plates that modify the properties of the flow ingested by the inlet. One plate may be oscillated to generate small perturbations in the flow. The other plate is held stationary to maintain a constant angle-of-attack. Using an idealized approach, design equations and performance maps for the new device were developed from the compressible flow relations. A two-dimensional CFD code was used to confirm the correctness of these results. The idealized approach was then used to design and evaluate a new gust generator for a 3.05-meter by 3.05-meter (10-foot by 10-foot) supersonic wind tunnel.

Performance of Single-Stage Turbine of Mark 25 Torpedo Power Plant with Two Nozzles and Three Rotor-Blade Designs

NASA Technical Reports Server (NTRS)

Schum, Harold J.; Whitney, Warren J.

1949-01-01

A single-stage modification of the turbine from a Mark 25 torpedo power plant was investigated to determine the performance with two nozzles and three rotor-blade designs. The performance was evaluated in terms of brake, rotor, and blade efficiencies at pressure ratios of 8, 15 (design), and 20. The blade efficiencies with the two nozzles are compared with those obtained with four other nozzles previously investigated with the same three rotor-blade designs. Blade efficiency with the cast nozzle of rectangular cross section (J) was higher than that with the circular reamed nozzle (K) at all speeds and pressure ratios with a rotor having a 0.45-inch 17 degree-inlet-angle blades. The efficiencies for both these nozzles were generally low compared with those of the four other nozzles previously investigated in combination with this rotor. At pressure ratios of 15 and 20, the blade efficiencies with nozzle K and the two rotors with 0.40-inch blades having different inlet angles were higher than with the four other nozzles, but the efficiency with nozzle J was generally low. Increasing the blade inlet angle from 17 degrees to 20 degrees had little effect on turbine performance, whereas changing the blade length from 0.40 to 0.45 inch had a marked effect. Although a slight correlation of efficiency with nozzle size was noted for the rotor with 0.45-inch 17 degree-inlet-angle blades, no such effect was discernible ,for the two rotors with 0.40-inch blades.Losses in the supersonic air stream resulting from the complex flow path in the small air passages are probably a large percentage of the total losses, and apparently the effects of changing nozzle size and shape within the limits investigated are of secondary importance.

Computer program for preliminary design analysis of axial-flow turbines

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1972-01-01

The program method is based on a mean-diameter flow analysis. Input design requirements include power or pressure ratio, flow, temperature, pressure, and speed. Turbine designs are generated for any specified number of stages and for any of three types of velocity diagrams (symmetrical, zero exit swirl, or impulse). Exit turning vanes can be included in the design. Program output includes inlet and exit annulus dimensions, exit temperature and pressure, total and static efficiencies, blading angles, and last-stage critical velocity ratios. The report presents the analysis method, a description of input and output with sample cases, and the program listing.

Experimental and numerical investigation of the iso-thermal flow characteristics within a cylindrical chamber with multiple planar-symmetric impinging jets

NASA Astrophysics Data System (ADS)

Long, Shen; Lau, Timothy C. W.; Chinnici, Alfonso; Tian, Zhao Feng; Dally, Bassam B.; Nathan, Graham J.

2017-10-01

We present a joint experimental and numerical study of the flow structure within a cylindrical chamber generated by planar-symmetric isothermal jets, under conditions of relevance to a wide range of practical applications, including the Hybrid Solar Receiver Combustor (HSRC) technology. The HSRC features a cavity with a coverable aperture to allow it to be operated as either a combustion chamber or a solar receiver, with multiple burners to direct a flame into the chamber and a heat exchanger that absorbs the heat from both energy sources. In this study, we assess the cases of two or four inlet jets (simulating the burners), configured in a planar-symmetric arrangement and aligned at an angle to the axis (αj) over the range of 0°-90°, at a constant inlet Reynolds number of ReD = 10 500. The jets were positioned in the same axial plane near the throat and interact with each other and the cavity walls. Measurements obtained with particle image velocimetry were used together with numerical modeling employing Reynolds-averaged Navier-Stokes methods to characterize the large-scale flow field within selected configurations of the device. The results reveal a significant dependence of the mean flow-field on αj and the number of inlet jets (Nj). Four different flow regimes with key distinctive features were identified within the range of αj and Nj considered here. It was also found that αj has a controlling influence on the extent of back-flow through the throat, the turbulence intensity, the flow stability, and the dominant recirculation zone, while Nj has a secondary influence on the turbulence intensity, the flow stability, and the transition between each flow regime.

Test data report, low speed wind tunnel tests of a full scale lift/cruise-fan inlet, with engine, at high angles of attack

NASA Technical Reports Server (NTRS)

Shain, W. M.

1978-01-01

A low speed wind tunnel test of a fixed lip inlet with engine, was performed. The inlet was close coupled to a Hamilton Standard 1.4 meter, variable pitch fan driven by a lycoming T55-L-11A engine. Tests were conducted with various combinations of inlet angle of attack freestream velocities, and fan airflows. Data were recorded to define the inlet airflow separation boundaries, performance characteristics, and fan blade stresses. The test model, installation, instrumentation, test, data reduction and final data are described.

Requirements for Large Eddy Simulation Computations of Variable-Speed Power Turbine Flows

NASA Technical Reports Server (NTRS)

Ameri, Ali A.

2016-01-01

Variable-speed power turbines (VSPTs) operate at low Reynolds numbers and with a wide range of incidence angles. Transition, separation, and the relevant physics leading to them are important to VSPT flow. Higher fidelity tools such as large eddy simulation (LES) may be needed to resolve the flow features necessary for accurate predictive capability and design of such turbines. A survey conducted for this report explores the requirements for such computations. The survey is limited to the simulation of two-dimensional flow cases and endwalls are not included. It suggests that a grid resolution necessary for this type of simulation to accurately represent the physics may be of the order of Delta(x)+=45, Delta(x)+ =2 and Delta(z)+=17. Various subgrid-scale (SGS) models have been used and except for the Smagorinsky model, all seem to perform well and in some instances the simulations worked well without SGS modeling. A method of specifying the inlet conditions such as synthetic eddy modeling (SEM) is necessary to correctly represent the inlet conditions.

Forced response analysis of an aerodynamically detuned supersonic turbomachine rotor

NASA Technical Reports Server (NTRS)

Hoyniak, D.; Fleeter, S.

1985-01-01

High performance aircraft-engine fan and compressor blades are vulnerable to aerodynamically forced vibrations generated by inlet flow distortions due to wakes from upstream blade and vane rows, atmospheric gusts, and maldistributions in inlet ducts. In this report, an analysis is developed to predict the flow-induced forced response of an aerodynamically detuned rotor operating in a supersonic flow with a subsonic axial component. The aerodynamic detuning is achieved by alternating the circumferential spacing of adjacent rotor blades. The total unsteady aerodynamic loading acting on the blading, as a result of the convection of the transverse gust past the airfoil cascade and the resulting motion of the cascade, is developed in terms of influence coefficients. This analysis is used to investigate the effect of aerodynamic detuning on the forced response of a 12-blade rotor, with Verdon's Cascade B flow geometry as a uniformly spaced baseline configuration. The results of this study indicate that, for forward traveling wave gust excitations, aerodynamic detuning is very beneficial, resulting in significantly decreased maximum-amplitude blade responses for many interblade phase angles.

Computational Fluid Dynamic Simulation of Flow in Abrasive Water Jet Machining

NASA Astrophysics Data System (ADS)

Venugopal, S.; Sathish, S.; Jothi Prakash, V. M.; Gopalakrishnan, T.

2017-03-01

Abrasive water jet cutting is one of the most recently developed non-traditional manufacturing technologies. In this machining, the abrasives are mixed with suspended liquid to form semi liquid mixture. The general nature of flow through the machining, results in fleeting wear of the nozzle which decrease the cutting performance. The inlet pressure of the abrasive water suspension has main effect on the major destruction characteristics of the inner surface of the nozzle. The aim of the project is to analyze the effect of inlet pressure on wall shear and exit kinetic energy. The analysis could be carried out by changing the taper angle of the nozzle, so as to obtain optimized process parameters for minimum nozzle wear. The two phase flow analysis would be carried by using computational fluid dynamics tool CFX. It is also used to analyze the flow characteristics of abrasive water jet machining on the inner surface of the nozzle. The availability of optimized process parameters of abrasive water jet machining (AWJM) is limited to water and experimental test can be cost prohibitive. In this case, Computational fluid dynamics analysis would provide better results.

Control of flow separation and mixing by aerodynamic excitation

NASA Technical Reports Server (NTRS)

Rice, Edward J.; Abbott, John M.

1990-01-01

The recent research in the control of shear flows using unsteady aerodynamic excitation conducted at the NASA Lewis Research Center is reviewed. The program is of a fundamental nature, concentrating on the physics of the unsteady aerodynamic processes. This field of research is a fairly new development with great promise in the areas of enhanced mixing and flow separation control. Enhanced mixing research includes influence of core turbulence, forced pairing of coherent structures, and saturation of mixing enhancement. Separation flow control studies included are for a two-dimensional diffuser, conical diffusers, and single airfoils. Ultimate applications include aircraft engine inlet flow control at high angle of attack, wide angle diffusers, highly loaded airfoils as in turbomachinery, and ejector/suppressor nozzles for the supersonic transport. An argument involving the Coanda Effect is made that all of the above mentioned application areas really only involve forms of shear layer mixing enhancement. The program also includes the development of practical excitation devices which might be used in aircraft applications.

Inlet-engine matching for SCAR including application of a bicone variable geometry inlet

NASA Technical Reports Server (NTRS)

Wasserbauer, J. F.; Gerstenmaier, W. H.

1978-01-01

Airflow characteristics of variable cycle engines (VCE) designed for Mach 2.32 can have transonic airflow requirements as high as 1.6 times the cruise airflow. This is a formidable requirement for conventional, high performance, axisymmetric, translating centerbody mixed compression inlets. An alternate inlet is defined, where the second cone of a two cone center body collapses to the initial cone angle to provide a large off-design airflow capability, and incorporates modest centerbody translation to minimize spillage drag. Estimates of transonic spillage drag are competitive with those of conventional translating centerbody inlets. The inlet's cruise performance exhibits very low bleed requirements with good recovery and high angle of attack capability.

Numerical analysis of the transient flow in a scroll refrigeration compressor

NASA Astrophysics Data System (ADS)

Sun, Shuaihui; Wu, Kai; Guo, Pengcheng; Luo, Xingqi

2017-08-01

In the present paper, the CFD technology is adopted to simulate the working process of a scroll refrigeration compressor with R22 as working fluid. The structural grids in the scroll compressor were updated continually during the solving process to cope with the movement boundaries of the fluid domain. The radial meshing clearance was 0.008 mm which was the same with that in the real prototype. The pressure, velocity and temperature distribution in chambers of compressor were computed. Also, the transient mass flux diagrams were calculated out. The results indicated that the pressure was asymmetrical in the two symmetrical suction chambers, because the suction port and passage were not absolutely symmetrical. The gradient of temperature was great in each working chamber due to leakage flow. Velocity vector distribution was asymmetrical in each pair of working chamber owing to the movement of orbiting scroll; the flow was complicated in the central working chamber. The movement of the orbiting scroll had different influence on the vortexes formation in each pair of compression chamber. The inlet and outlet mass flux fluctuated with the crank angle obviously. Because of the ‘cut-off’ of the refrigeration fluid in the suction chamber when the crank angle was larger than 220°, the inlet mass flux decreased remarkably. Finally, some useful advices were given to improve the performance of the scroll refrigeration compressor.

Simulation of aerosolized oil droplets capture in a range hood exhaust using coupled CFD-population balance method

NASA Astrophysics Data System (ADS)

Liu, Shuyuan; Zhang, Yong; Feng, Yu; Shi, Changbin; Cao, Yong; Yuan, Wei

2018-02-01

A coupled population balance sectional method (PBSM) coupled with computational fluid dynamics (CFD) is presented to simulate the capture of aerosolized oil droplets (AODs) in a range hood exhaust. The homogeneous nucleation and coagulation processes are modeled and simulated with this CFD-PBSM method. With the design angle, α of the range hood exhaust varying from 60° to 30°, the AODs capture increases meanwhile the pressure drop between the inlet and the outlet of the range hood also increases from 8.38Pa to 175.75Pa. The increasing inlet flow velocities also result in less AODs capture although the total suction increases due to higher flow rates to the range hood. Therefore, the CFD-PBSM method provides an insight into the formation and capture of AODs as well as their impact on the operation and design of the range hood exhaust.

Aerodynamics of powered missile separation from F/A-18 aircraft

NASA Technical Reports Server (NTRS)

Ahmad, J. U.; Shanks, S. P.; Buning, P. G.

1993-01-01

A 3D dynamic 'chimera' algorithm that solves the thin-layer Navier-Stokes equations over multiple moving bodies was modified to numerically simulate the aerodynamics, missile dynamics, and missile plume interactions of a missile separating from a generic wing and from an F/A-18 aircraft in transonic flow. The missile is mounted below the wing for missile separation from the wing and on the F/A-18 fuselage at the engine inlet side for missile separation from aircraft. Static and powered missile separation cases are considered to examine the influence of the missile and plume on the wing and F/A-18 fuselage and engine inlet. The aircraft and missile are at two degrees angle of attack, Reynolds number of 10 million, freestream Mach number of 1.05 and plume Mach number of 3.0. The computational results show the details of the flow field.

Hydrocarbon fluid, ejector refrigeration system

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

Kowalski, G.J.; Foster, A.R.

1993-08-31

A refrigeration system is described comprising: a vapor ejector cycle including a working fluid having a property such that entropy of the working fluid when in a saturated vapor state decreases as pressure decreases, the vapor ejector cycle comprising: a condenser located on a common fluid flow path; a diverter located downstream from the condenser for diverting the working fluid into a primary fluid flow path and a secondary fluid flow path parallel to the primary fluid flow path; an evaporator located on the secondary fluid flow path; an expansion device located on the secondary fluid flow path upstream ofmore » the evaporator; a boiler located on the primary fluid flow path parallel to the evaporator for boiling the working fluid, the boiler comprising an axially extending core region having a substantially constant cross sectional area and a porous capillary region surrounding the core region, the core region extending a length sufficient to produce a near sonic velocity saturated vapor; and an ejector having an outlet in fluid communication with the inlet of the condenser and an inlet in fluid communication with the outlet of the evaporator and the outlet of the boiler and in which the flows of the working fluid from the evaporator and the boiler are mixed and the pressure of the working fluid is increased to at least the pressure of the condenser, the ejector inlet, located downstream from the axially extending core region, including a primary nozzle located sufficiently close to the outlet of the boiler to minimize a pressure drop between the boiler and the primary nozzle, the primary nozzle of the ejector including a converging section having an included angle and length preselected to receive the working fluid from the boiler as a near sonic velocity saturated vapor.« less

Hydrodynamic and thermal mechanisms of filtration combustion inclinational instability based on non-uniform distribution of initial preheating temperature

NASA Astrophysics Data System (ADS)

Xia, Yongfang; Shi, Junrui; Xu, Youning; Ma, Rui

2018-03-01

Filtration combustion (FC) is one style of porous media combustion with inert matrix, in which the combustion wave front propagates, only downstream or reciprocally. In this paper, we investigate the FC flame front inclinational instability of lean methane/air mixtures flowing through a packed bed as a combustion wave front perturbation of the initial preheating temperature non-uniformity is assumed. The predicted results show that the growth rate of the flame front inclinational angle is proportional to the magnitude of the initial preheating temperature difference. Additionally, depending on gas inlet gas velocity and equivalence ratio, it is demonstrated that increase of gas inlet gas velocity accelerates the FC wave front deformation, and the inclinational instability evolves faster at lower equivalence ratio. The development of the flame front inclinational angle may be regarded as a two-staged evolution, which includes rapid increase, and approaching maximum value of inclinational angle due to the quasi-steady condition of the combustion system. The hydrodynamic and thermal mechanisms of the FC inclinational instability are analyzed. Consequently, the local propagation velocity of the FC wave front is non-uniform to result in the development of inclinational angle at the first stage of rapid increase.

Full-Authority Fault-Tolerant Electronic Engine Control Systems for Variable Cycle Engines.

DTIC Science & Technology

1981-12-01

Geometry or Fuel Flow Scheduled as a Function of Engine State, i.e. FIGV = f( N1 C2 ) Closed Loop - Geometry or Fuel Flow Modulated To Maintain an Engine...Low Pressure Turbine Inlet Area (A41) Closed Loop (Integral) N2, T22 Core Stream Exhaust Nozzle Area (AJE) Closed Loop (Integral) N1 , T2 Duct Stream...to remain at the breakpoint value while low rotor speed reference ( N1 reference) is scheduled to decrease as a function of power lever angle (PLA), to

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.

SSME Turbopump Turbine Computations

NASA Technical Reports Server (NTRS)

Jorgenson, P. G. E.

1985-01-01

A two-dimensional viscous code was developed to be used in the prediction of the flow in the SSME high-pressure turbopump blade passages. The rotor viscous code (RVC) employs a four-step Runge-Kutta scheme to solve the two-dimensional, thin-layer Navier-Stokes equations. The Baldwin-Lomax eddy-viscosity model is used for these turbulent flow calculations. A viable method was developed to use the relative exit conditions from an upstream blade row as the inlet conditions to the next blade row. The blade loading diagrams are compared with the meridional values obtained from an in-house quasithree-dimensional inviscid code. Periodic boundary conditions are imposed on a body-fitted C-grid computed by using the GRAPE GRids about Airfoils using Poisson's Equation (GRAPE) code. Total pressure, total temperature, and flow angle are specified at the inlet. The upstream-running Riemann invariant is extrapolated from the interior. Static pressure is specified at the exit such that mass flow is conserved from blade row to blade row, and the conservative variables are extrapolated from the interior. For viscous flows the noslip condition is imposed at the wall. The normal momentum equation gives the pressure at the wall. The density at the wall is obtained from the wall total temperature.

Aerodynamic Design of Axial-flow Compressors. Volume 2

NASA Technical Reports Server (NTRS)

1956-01-01

Available experimental two-dimensional-cascade data for conventional compressor blade sections are correlated. The two-dimensional cascade and some of the principal aerodynamic factors involved in its operation are first briefly described. Then the data are analyzed by examining the variation of cascade performance at a reference incidence angle in the region of minimum loss. Variations of reference incidence angle, total-pressure loss, and deviation angle with cascade geometry, inlet Mach number, and Reynolds number are investigated. From the analysis and the correlations of the available data, rules and relations are evolved for the prediction of the magnitude of the reference total-pressure loss and the reference deviation and incidence angles for conventional blade profiles. These relations are developed in simplified forms readily applicable to compressor design procedures.

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.

Investigation of coaxial jet noise and inlet choking using an F-111A airplane

NASA Technical Reports Server (NTRS)

Putnam, T. W.

1973-01-01

Measurements of engine noise generated by an F-111A airplane positioned on a thrustmeasuring platform were made at angles of 0 deg to 160 deg from the aircraft heading. Sound power levels, power spectra, and directivity patterns are presented for jet exit velocities between 260 feet per second and 2400 feet per second. The test results indicate that the total acoustic power was proportional to the eighth power of the core jet velocity for core exhaust velocities greater than 300 meters per second (985 feet per second) and that little or no mixing of the core and fan streams occurred. The maximum sideline noise was most accurately predicted by using the average jet velocity for velocities above 300 meters per second (985 feet per second). The acoustic power spectrum was essentially the same for the single jet flow of afterburner operation and the coaxial flow of the nonafterburning condition. By varying the inlet geometry and cowl position, reductions in the sound pressure level of the blade passing frequency on the order of 15 decibels to 25 decibels were observed for inlet Mach numbers of 0.8 to 0.9.

Mach 6 flow field surveys beneath the forebody of an airbreathing missile

NASA Technical Reports Server (NTRS)

Johnson, P. J.; Hunt, J. L.

1986-01-01

Wall static, local stream static, and pitot pressure surveys were made on the windward side of a hypersonic airbreathing missile at full-scale length Reynolds numbers. In the inviscid part of the flow field, the experimental massflow ratios agreed with trends predicted by a three-dimensional method-of-characteristics solution. At a longitudinal station 3.5 diameters downstrea of the nose, the boundary layer was transitional or turbulent at zero incidence but became laminar as the angle of attack increased. The bell-shaped distribution of the boundary layer across the width of the body affected the mass flow distribution out to the bow shock and decreased the mass flow available the engine inlet.

Laser velocimeter measurements of the flow downstream of the Space Shuttle Main Engine high pressure oxidizer turbopump first-stage turbine nozzle

NASA Technical Reports Server (NTRS)

Ferguson, T. V.; Havskjold, G. L.; Rojas, L.

1988-01-01

A laser two-focus velocimeter was used in an open-loop water test facility in order to map the flowfield downstream of the SSME's high-pressure oxidizer turbopump first-stage turbine nozzle; attention was given to the effects of the upstream strut-downstream nozzle configuration on the flow at the rotor inlet, in order to estimate dynamic loads on the first-stage rotor blades. Velocity and flow angles were plotted as a function of circumferential position, and were found to clearly display the periodic behavior of the wake flow field. The influence of the upstream centerbody-supporting struts on the vane nozzle wake pattern was evident.

A forebody design technique for highly integrated bottom-mounted scramjets with application to a hypersonic research airplane

NASA Technical Reports Server (NTRS)

Edwards, C. L. W.

1974-01-01

An inviscid technique for designing forebodies which produce uniformly precompressed flows at the inlet entrance for bottom-mounted scramjets has been developed so that geometric constraints resulting from design trade-offs can be effectively evaluated. The flow fields resulting from several forebody designs generated in support of a hypersonic research airplane conceptual design study have been analyzed in detail with three-dimensional characteristics calculations to verify the uniform flow conditions. For the designs analyzed, uniform flow is maintained over a wide range of flight conditions (Mach number equals 4 to 10; angle of attack equals 6 deg to 10 deg) corresponding to scramjet operation flight envelope of the research airplane.

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.

Design Evolution and Performance Characterization of the GTX Air-Breathing Launch Vehicle Inlet

NASA Technical Reports Server (NTRS)

DeBonis, J. R.; Steffen, C. J., Jr.; Rice, T.; Trefny, C. J.

2002-01-01

The design and analysis of a second version of the inlet for the GTX rocket-based combine-cycle launch vehicle is discussed. The previous design did not achieve its predicted performance levels due to excessive turning of low-momentum comer flows and local over-contraction due to asymmetric end-walls. This design attempts to remove these problems by reducing the spike half-angle to 10- from 12-degrees and by implementing true plane of symmetry end-walls. Axisymmetric Reynolds-Averaged Navier-Stokes simulations using both perfect gas and real gas, finite rate chemistry, assumptions were performed to aid in the design process and to create a comprehensive database of inlet performance. The inlet design, which operates over the entire air-breathing Mach number range from 0 to 12, and the performance database are presented. The performance database, for use in cycle analysis, includes predictions of mass capture, pressure recovery, throat Mach number, drag force, and heat load, for the entire Mach range. Results of the computations are compared with experimental data to validate the performance database.

Radial inlet guide vanes for a combustor

DOEpatents

Zuo, Baifang; Simons, Derrick; York, William; Ziminsky, Willy S

2013-02-12

A combustor may include an interior flow path therethrough, a number of fuel nozzles in communication with the interior flow path, and an inlet guide vane system positioned about the interior flow path to create a swirled flow therein. The inlet guide vane system may include a number of windows positioned circumferentially around the fuel nozzles. The inlet guide vane system may also include a number of inlet guide vanes positioned circumferentially around the fuel nozzles and adjacent to the windows to create a swirled flow within the interior flow path.

Performance Comparison at Mach Numbers 1.8 and 2.0 of Full Scale and Quarter Scale Translating-Spike Inlets

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Dryer, M.; Hearth, D. P.

1957-01-01

The performance of a full-scale translating-spike inlet was obtained at Mach numbers of 1.8 and 2.0 and at angles of attach from 0 deg to 6 deg. Comparisons were made between the full-scale production inlet configuration and a geometrically similar quarter-scale model. The inlet pressure-recovery, cowl pressure-distribution, and compressor-face distortion characteristics of the full-scale inlet agreed fairly well with the quarter-scale results. In addition, the results indicated that bleeding around the periphery ahead of the compressor-face station improved pressure recovery and compressor-face distortion, especially at angle of attack.

Effect of cationic monomer content on polyacrylamide copolymers by frit-inlet asymmetrical flow field-flow fractionation/multi-angle light scattering.

PubMed

Lee, Hyejin; Kim, Jin Yong; Choi, Woonjin; Moon, Myeong Hee

2017-06-23

In this study, ultrahigh-molecular-weight (MW) (>10 7 Da) cationic polyacrylamides (C-PAMs), which are water-soluble polymers used in waste water treatment, were characterized using frit-inlet asymmetrical flow field-flow fractionation coupled with multi-angle light scattering and differential refractive detection. C-PAMs copolymerized with acryloxyethyltrimethyl ammonium chloride (DAC) were prepared by varying the feed amount of cationic monomer, polymerization method (solution vs. emulsion), and degree of branching. The MW of the copolymers prepared using emulsion polymerization (10 7 -10 9 Da) was generally larger than that of copolymers prepared using solution polymerization (4×10 7 -10 8 Da). When the amount of cationic monomer was increased from 10 to 55mol% in solution polymerization, hydrophobic contraction of the core induced formation of more compact C-PAMs. The copolymers prepared using emulsion polymerization formed highly aggregated or supercoil structures owing to increased intermolecular hydrophobic interaction when less cationic monomer was used. However, the MW decreased with increased cationic group content. In addition, C-PAMs larger than ∼10 8 Da prepared using the emulsion method were separated by steric/hyperlayer elution mode while those in the 10 7 -10 8 Da range were analyzed in either normal or steric/hyperlayer mode depending on the decay patterns of field programming. Moreover, branched copolymers were found to be resolved with different elution modes under the same field decay pattern depending on the degree of branching: steric/hyperlayer for low-branching and normal for high-branching C-PAMs. Copyright © 2017 Elsevier B.V. All rights reserved.

Investigation of normal shock inlets for highly maneuverable aircraft

NASA Technical Reports Server (NTRS)

Martin, A. W.

1977-01-01

Concepts are investigated for obtaining both low cowl drag and good inlet performance at high angles of attack. The effect of a canard on inlet performance for a kidney shaped inlet in each of two vertical locations is discussed along with a sharp lip two dimensional inlet on a canardless forebody.

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.

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.

Experimental investigation of the inlet detector configuration variation in the flow field at Mach 1.9

NASA Technical Reports Server (NTRS)

Hwang, Kyu C.; Tiwari, Surrendra N.; Miley, Stanley J.

1995-01-01

In recent years, active research has been conducted to study the technological feasibility of supersonic laminar flow control on the wing of the High Speed Civil Transport (HSCT). For this study, the F-16XL has been chosen due to its highly swept crank wing planform that closely resembles the HSCT configurations. During flights, it is discovered that the shock wave generated from the aircraft inlet introduces disturbances on the wing where the data acquisition is conducted. The flow field about a supersonic inlet is characterized by a complex three dimensional pattern of shock waves generated by the geometrical configuration of a deflector and a cowl lip. Hence, in this study, experimental method is employed to investigate the effects of the variation of deflector configuration on the flow field, and consequently, the possibility of diverting the incoming shock-disturbances away from the test section. In the present experiments, a model composed of a simple circular tube with a triangular deflector is designed to study the deflector length and the deflector base width variation in the flow field. Experimental results indicate that the lowest external pressure ratio is observed at the junction where the deflector lip and the inlet cowl lip merge. Also, it is noted that the external pressure ratio, the internal pressure ratio, the coefficient of spillage drag, and the shock standoff distance decrease as the deflector length increases. In addition, the Redefined Total Pressure Recovery Ratio (RTPRR) increases with an increase in the deflector length. Results from the study of the effect of the deflector's base width variation on the flow field indicate that the lowest external pressure ratio is observed at the junction between the inlet cowl lip and the deflector lip. As the base width of the deflector increases, the external pressure ratio at 0 rotation increases, whereas the external pressure ratio at 180 rotation decreases. In addition, the internal pressure ratio and the coefficient of spillage drag decrease as the base width of the deflector increases. However, RTPRR and shock standoff distance increase as the base width increases. In conclusion, as deflector dimensions vary, distinctive patterns in the pressure variation around the inlet deflector are observed. With an increase in the deflector length and base width, the magnitude of shock-disturbances are weakened due to a decrease in the external pressure ratio. Also, as the deflector length and base width increase, a smaller bow shock angle is formed. Therefore, the inlet shock wave formation would be significantly altered, and consequently, shock disturbances on the wing test section can be avoided through appropriately designing the deflector.

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.

Performance Improvement of a Return Channel in a Multistage Centrifugal Compressor Using Multiobjective Optimization.

PubMed

Nishida, Yoshifumi; Kobayashi, Hiromi; Nishida, Hideo; Sugimura, Kazuyuki

2013-05-01

The effect of the design parameters of a return channel on the performance of a multistage centrifugal compressor was numerically investigated, and the shape of the return channel was optimized using a multiobjective optimization method based on a genetic algorithm to improve the performance of the centrifugal compressor. The results of sensitivity analysis using Latin hypercube sampling suggested that the inlet-to-outlet area ratio of the return vane affected the total pressure loss in the return channel, and that the inlet-to-outlet radius ratio of the return vane affected the outlet flow angle from the return vane. Moreover, this analysis suggested that the number of return vanes affected both the loss and the flow angle at the outlet. As a result of optimization, the number of return vane was increased from 14 to 22 and the area ratio was decreased from 0.71 to 0.66. The radius ratio was also decreased from 2.1 to 2.0. Performance tests on a centrifugal compressor with two return channels (the original design and optimized design) were carried out using two-stage test apparatus. The measured flow distribution exhibited a swirl flow in the center region and a reversed swirl flow near the hub and shroud sides. The exit flow of the optimized design was more uniform than that of the original design. For the optimized design, the overall two-stage efficiency and pressure coefficient were increased by 0.7% and 1.5%, respectively. Moreover, the second-stage efficiency and pressure coefficient were respectively increased by 1.0% and 3.2%. It is considered that the increase in the second-stage efficiency was caused by the increased uniformity of the flow, and the rise in the pressure coefficient was caused by a decrease in the residual swirl flow. It was thus concluded from the numerical and experimental results that the optimized return channel improved the performance of the multistage centrifugal compressor.

L.D.V. measurements of unsteady flow fields in radial turbine

NASA Astrophysics Data System (ADS)

Tabakoff, W.; Pasin, M.

1992-07-01

Detailed measurements of an unsteady flow field within the inlet guide vanes (IGV) and the rotor of a radial inflow turbine were performed using a three component Laser Doppler Velocimeter (LDV) system together with a rotary encoder. The mean velocity, the flow angle and the turbulence contours for IGV passages are presented at four blade-to-blade planes for different rotor positions to give three dimensional, unsteady behavior of the IGV flow field. These results are compared with the measurements obtained in the same passage in the absence of the rotor. The flow field of the IGV passage was found to be affected by the presence of the rotor. The ratio of the tangential normal stresses to the radial normal stresses at the exit of the IGV was found to be more than doubled when compared to the case without the rotor. The rotor flow field measurements are presented as relative mean velocity and turbulence stress contours at various cross section planes throughout the rotor. The cross flow and turbulence stress levels were found to be influenced by the incidence angle. Transportation of the high turbulence fluid by the cross flow was observed downstream in the rotor blade passages.

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.

DESIGN OF TWO-DIMENSIONAL SUPERSONIC TURBINE ROTOR BLADES WITH BOUNDARY-LAYER CORRECTION

NASA Technical Reports Server (NTRS)

Goldman, L. J.

1994-01-01

A computer program has been developed for the design of supersonic rotor blades where losses are accounted for by correcting the ideal blade geometry for boundary layer displacement thickness. The ideal blade passage is designed by the method of characteristics and is based on establishing vortex flow within the passage. Boundary-layer parameters (displacement and momentum thicknesses) are calculated for the ideal passage, and the final blade geometry is obtained by adding the displacement thicknesses to the ideal nozzle coordinates. The boundary-layer parameters are also used to calculate the aftermixing conditions downstream of the rotor blades assuming the flow mixes to a uniform state. The computer program input consists essentially of the rotor inlet and outlet Mach numbers, upper- and lower-surface Mach numbers, inlet flow angle, specific heat ratio, and total flow conditions. The program gas properties are set up for air. Additional gases require changes to be made to the program. The computer output consists of the corrected rotor blade coordinates, the principal boundary-layer parameters, and the aftermixing conditions. This program is written in FORTRAN IV for batch execution and has been implemented on an IBM 7094. This program was developed in 1971.

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.

Performance and surge limits of a TF30-P-3 turbofan engine/axisymmetric mixed-compression inlet propulsion system at Mach 2.5

NASA Technical Reports Server (NTRS)

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

1985-01-01

Steady-state performance and inlet-engine compatibility were investigated with a low-bleed inlet. The inlet had minimum internal contraction, consistent with high total pressure recovery and low cowl drag. The inlet-engine combination displayed good performance with only about 2% of inlet performance bleed. The inlet-engine combination had 5.58 deg angle-of-attack capability with 6% bleed.

Advanced Technology Inlet Design, NRA 8-21 Cycle II: DRACO Flowpath Hypersonic Inlet Design

NASA Technical Reports Server (NTRS)

Sanders, Bobby W.; Weir, Lois J.

1999-01-01

The report outlines work performed in support of the flowpath development for the DRACO engine program. The design process initiated to develop a hypersonic axisymmetric inlet for a Mach 6 rocket-based combined cycle (RBCC) engine is discussed. Various design parametrics were investigated, including design shock-on-lip Mach number, cone angle, throat Mach number, throat angle. length of distributed compression, and subsonic diffuser contours. Conceptual mechanical designs consistent with installation into the D-21 vehicle were developed. Additionally, program planning for an intensive inlet development program to support a Critical Design Review in three years was performed. This development program included both analytical and experimental elements and support for a flight-capable inlet mechanical design.

Supersonic Inlet Flow Control Using Localized Arc Filament Plasma Actuators

DTIC Science & Technology

2011-05-10

Jon Schmisseur and by the Air Vehicle Directorate of the Air Force Research Laboratory is greatly appreciated. 4 Table of Contents Abstract...of interest to the U.S. Air Force and could pose significant problems depending on the specific application. This study has undertaken to investigate...experiments motivated the design of a new, larger, more flexible facility that utilize a Variable Angle Wedge to generate the impinging shock wave for

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.

Complementary Aerodynamic Performance Datasets for Variable Speed Power Turbine Blade Section from Two Independent Transonic Turbine Cascades

NASA Technical Reports Server (NTRS)

Flegel, Ashlie B.; Welch, Gerard E.; Giel, Paul W.; Ames, Forrest E.; Long, Jonathon A.

2015-01-01

Two independent experimental studies were conducted in linear cascades on a scaled, two-dimensional mid-span section of a representative Variable Speed Power Turbine (VSPT) blade. The purpose of these studies was to assess the aerodynamic performance of the VSPT blade over large Reynolds number and incidence angle ranges. The influence of inlet turbulence intensity was also investigated. The tests were carried out in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility and at the University of North Dakota (UND) High Speed Compressible Flow Wind Tunnel Facility. A large database was developed by acquiring total pressure and exit angle surveys and blade loading data for ten incidence angles ranging from +15.8deg to -51.0deg. Data were acquired over six flow conditions with exit isentropic Reynolds number ranging from 0.05×106 to 2.12×106 and at exit Mach numbers of 0.72 (design) and 0.35. Flow conditions were examined within the respective facility constraints. The survey data were integrated to determine average exit total-pressure and flow angle. UND also acquired blade surface heat transfer data at two flow conditions across the entire incidence angle range aimed at quantifying transitional flow behavior on the blade. Comparisons of the aerodynamic datasets were made for three "match point" conditions. The blade loading data at the match point conditions show good agreement between the facilities. This report shows comparisons of other data and highlights the unique contributions of the two facilities. The datasets are being used to advance understanding of the aerodynamic challenges associated with maintaining efficient power turbine operation over a wide shaft-speed range.

Test data report: Low speed wind tunnel tests of a full scale, fixed geometry inlet, with engine, at high angles of attack

NASA Technical Reports Server (NTRS)

Shain, W. M.

1976-01-01

A full scale inlet test was to be done in the NASA-ARC 40' X 80' WT to demonstrate satisfactory inlet performance at high angles of attack. The inlet was designed to match a Hamilton-Standard 55 inch, variable pitch fan, driven by a Lycoming T55-L-11A gas generator. The test was installed in the wind tunnel on two separate occasions, but mechanical failures in the fan drive gear box early in each period terminated testing. A detailed description is included of the Model, installation, instrumentation and data reduction procedures.

Static internal performance of a single-engine onaxisymmetric-nozzle vaned-thrust-reverser design with thrust modulation capabilities

NASA Technical Reports Server (NTRS)

Leavitt, L. D.; Burley, J. R., II

1985-01-01

An investigation has been conducted at wind-off conditions in the stati-test facility of the Langley 16-Foot Transonic Tunnel. The tests were conducted on a single-engine reverser configuration with partial and full reverse-thrust modulation capabilities. The reverser design had four ports with equal areas. These ports were angled outboard 30 deg from the vertical impart of a splay angle to the reverse exhaust flow. This splaying of reverser flow was intended to prevent impingement of exhaust flow on empennage surfaces and to help avoid inlet reingestion of exhaust gas when the reverser is integrated into an actual airplane configuration. External vane boxes were located directly over each of the four ports to provide variation of reverser efflux angle from 140 deg to 26 deg (measured forward from the horizontal reference axis). The reverser model was tested with both a butterfly-type inner door and an internal slider door to provide area control for each individual port. In addition, main nozzle throat area and vector angle were varied to examine various methods of modulating thrust levels. Other model variables included vane box configuration (four or six vanes per box), orientation of external vane boxes with respect to internal port walls (splay angle shims), and vane box sideplates. Nozzle pressure ratio was varied from 2.0 approximately 7.0.

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.

The Charging Process in a High-speed, Single-cylinder, Four-stroke Engine

NASA Technical Reports Server (NTRS)

Reynolds, Blake; Schecter, Harry; Taylor, E S

1939-01-01

Experimental measurements and theoretical calculations were made on an aircraft-type, single cylinder engine, in order to determine the physical nature of the inlet process, especially at high piston speeds. The engine was run at speeds from 1,500 to 2,600 r.p.m. (mean piston speeds of 1,370 to 2,380 feet per minute). Measurements were made of the cylinder pressure during the inlet stroke and of the power output and volumetric efficiency. Measurements were also made, with the engine not running, to determine the resistance and mass of air in the inlet valve port at various crank angles. Results of analysis indicate that mass has an appreciable effect, but friction plays the major part in restricting flow. The observed fact that the volumetric efficiency is considerably less than 100 percent is attributed to thermal effects. An estimate was made of the magnitude of these effects in the present case, and their general nature is discussed.

Numerical Optimization of converging diverging miniature cavitating nozzles

NASA Astrophysics Data System (ADS)

Chavan, Kanchan; Bhingole, B.; Raut, J.; Pandit, A. B.

2015-12-01

The work focuses on the numerical optimization of converging diverging cavitating nozzles through nozzle dimensions and wall shape. The objective is to develop design rules for the geometry of cavitating nozzles for desired end-use. Two main aspects of nozzle design which affects the cavitation have been studied i.e. end dimensions of the geometry (i.e. angle and/or curvature of the inlet, outlet and the throat and the lengths of the converging and diverging sections) and wall curvatures(concave or convex). Angle of convergence at the inlet was found to control the cavity growth whereas angle of divergence of the exit controls the collapse of cavity. CFD simulations were carried out for the straight line converging and diverging sections by varying converging and diverging angles to study its effect on the collapse pressure generated by the cavity. Optimized geometry configurations were obtained on the basis of maximum Cavitational Efficacy Ratio (CER)i.e. cavity collapse pressure generated for a given permanent pressure drop across the system. With increasing capabilities in machining and fabrication, it is possible to exploit the effect of wall curvature to create nozzles with further increase in the CER. Effect of wall curvature has been studied for the straight, concave and convex shapes. Curvature has been varied and effect of concave and convex wall curvatures vis-à-vis straight walls studied for fixed converging and diverging angles.It is concluded that concave converging-diverging nozzles with converging angle of 20° and diverging angle of 5° with the radius of curvature 0.03 m and 0.1530 m respectively gives maximum CER. Preliminary experiments using optimized geometry are indicating similar trends and are currently being carried out. Refinements of the CFD technique using two phase flow simulations are planned.

Inlet flow distortion in turbomachinery. I - Comparison of theory and experiment in a transonic fan stage. II - A parameter study

NASA Technical Reports Server (NTRS)

Seidel, B. S.; Matwey, M. D.; Adamczyk, J. J.

1980-01-01

In the present paper, a semi-actuator-disk theory is reviewed that was developed previously for the distorted inflow to a single-stage axial-flow compressor. Flow distortion occurs far upstream; it may be a distortion in stagnation temperature, stagnation pressure, or both. Losses, quasi-steady deviation angles, and reference incidence correlations are included in the analysis, and both subsonic and transonic relative Mach numbers are considered. The theory is compared with measurements made in a transonic fan stage, and a parameter study is carried out to determine the influence of solidity on the attenuation of distortions in stagnation pressure and stagnation temperature.

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.

Hover and wind-tunnel testing of shrouded rotors for improved micro air vehicle design

NASA Astrophysics Data System (ADS)

Pereira, Jason L.

The shrouded-rotor configuration has emerged as the most popular choice for rotary-wing Micro Air Vehicles (MAVs), because of the inherent safety of the design and the potential for significant performance improvements. However, traditional design philosophies based on experience with large-scale ducted propellers may not apply to the low-Reynolds-number (˜20,000) regime in which MAVs operate. An experimental investigation of the effects of varying the shroud profile shape on the performance of MAV-scale shrouded rotors has therefore been conducted. Hover tests were performed on seventeen models with a nominal rotor diameter of 16 cm (6.3 in) and various values of diffuser expansion angle, diffuser length, inlet lip radius and blade tip clearance, at various rotor collective angles. Compared to the baseline open rotor, the shrouded rotors showed increases in thrust by up to 94%, at the same power consumption, or reductions in power by up to 62% at the same thrust. These improvements surpass those predicted by momentum theory, due to the additional effect of the shrouds in reducing the non-ideal power losses of the rotor. Increasing the lip radius and decreasing the blade tip clearance caused performance to improve, while optimal values of diffuser angle and length were found to be 10 and 50% of the shroud throat diameter, respectively. With the exception of the lip radius, the effects of changing any of the shrouded-rotor parameters on performance became more pronounced as the values of the other parameters were changed to degrade performance. Measurements were also made of the wake velocity profiles and the shroud surface pressure distributions. The uniformity of the wake was improved by the presence of the shrouds and by decreasing the blade tip clearance, resulting in lower induced power losses. For high net shroud thrust, a favorable pressure distribution over the inlet was seen to be more important than in the diffuser. Strong suction pressures were observed above the blade-passage region on the inlet surface; taking advantage of this phenomenon could enable further increases in thrust. However, trade studies showed that, for a given overall aircraft size limitation, and ignoring considerations of the safety benefits of a shroud, a larger-diameter open rotor is more likely to give better performance than a smaller-diameter shrouded rotor. The open rotor and a single shrouded-rotor model were subsequently tested at a single collective in translational flight, at angles of attack from 0° (axial flow) to 90° (edgewise flow), and at various advance ratios. In axial flow, the net thrust and the power consumption of the shrouded rotor were lower than those of the open rotor. In edgewise flow, the shrouded rotor produced greater thrust than the open rotor, while consuming less power. Measurements of the shroud surface pressure distributions illustrated the extreme longitudinal asymmetry of the flow around the shroud, with consequent pitch moments much greater than those exerted on the open rotor. Except at low airspeeds and high angles of attack, the static pressure in the wake did not reach ambient atmospheric values at the diffuser exit plane; this challenges the validity of the fundamental assumption of the simple-momentum-theory flow model for short-chord shrouds in translational flight.

Application of the CSCM method to the design of wedge cavities. [Conservative Supra Characteristic Method

NASA Technical Reports Server (NTRS)

Venkatapathy, Ethiraj; Nystrom, G. A.; Bardina, J.; Lombard, C. K.

1987-01-01

This paper describes the application of the conservative supra characteristic method (CSCM) to predict the flow around two-dimensional slot injection cooled cavities in hypersonic flow. Seven different numerical solutions are presented that model three different experimental designs. The calculations manifest outer flow conditions including the effects of nozzle/lip geometry, angle of attack, nozzle inlet conditions, boundary and shear layer growth and turbulance on the surrounding flow. The calculations were performed for analysis prior to wind tunnel testing for sensitivity studies early in the design process. Qualitative and quantitative understanding of the flows for each of the cavity designs and design recommendations are provided. The present paper demonstrates the ability of numerical schemes, such as the CSCM method, to play a significant role in the design process.

Sheathless focusing and separation of microparticles using tilted angle travelling surface acoustic waves.

PubMed

Ahmed, Husnain; Destgeer, Ghulam; Park, Jinsoo; Afzal, Muhammad; Sung, Hyung Jin

2018-06-18

The sheathless focusing and separation of microparticles is an important pre-processing step in various biochemical assays in which enriched sample isolation is critical. Most previous microfluidic particle separation techniques have used a sheath flow to achieve efficient sample focusing. The sheath flow diluted the analyte, and required additional microchannels and accurate flow control. We demonstrated a tilted angle travelling surface acoustic wave (taTSAW)-based sheathless focusing and separation of particles in a continuous flow. The proposed device consisted of a piezoelectric substrate with a pair of interdigitated transducers (IDTs) deposited at two different angles relative to the flow direction. A Y-shaped polydimethylsiloxane (PDMS) microchannel having one inlet and two outlet ports was positioned on top of the IDTs such that the acoustic energy coupling into the fluid was maximized and wave attenuation by the PDMS walls was minimized. The two IDTs independently produced high-frequency taTSAWs, which propagated at ±30° with respect to the flow direction and imparted a direct acoustic radiation force onto the target particles. A sample mixture containing 4.8 and 3.2 µm particles was focused and then separated by the actuation of the IDTs at 194 and 136 MHz frequencies, respectively, without using an additional sheath flow. The proposed taTSAW-based particle separation device offered a high purity > 99% at the both outlets over a wide range of flow speeds (up to 83.3 mm/s).

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.

Experimental investigation on discharge coefficient for a combined broad crested weir-box culvert structure

NASA Astrophysics Data System (ADS)

Guven, A.; Hassan, M.; Sabir, Shahin

2013-09-01

This study investigates the hydraulic characteristics of simultaneous flow over broad crested weir and through box (square) culverts experimentally. The variation of discharge coefficient (Cd) of the combined structure with various affective parameters such as upstream head, length of culvert, culvert inlet shape, culvert internal dimension, weir crest height, weir side slope angle, and weir width was analyzed. For this purpose 12 glass models of combined broad crested weirs and box culverts were manufactured and tested in a laboratory flume of 12 m. Discharge coefficient predicting equations were developed as a function of the dimensionless terms. The overall results showed that Cd increases as the ratio of the total head of water above the weir crest to the height of the weir crest H/P increases for all the models and for each flow state (weir and combined). Cd values increased as the head increased for all the models tested as culvert flow only, and also with decreasing of the angle between crest of the weir with the sides.

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.

Characterization of ultrahigh-molecular weight cationic polyacrylamide using frit-inlet asymmetrical flow field-flow fractionation and multi-angle light scattering.

PubMed

Woo, Sohee; Lee, Ju Yong; Choi, Woonjin; Moon, Myeong Hee

2016-01-15

In this study, frit inlet asymmetrical flow field-flow fractionation (FlFFF) with multi-angle light scattering (MALS) and differential refractive index (DRI) detection is utilized for size separation, determination of molecular weight (MW), and conformation of ultrahigh-MW (10(7)-10(9) g/mol) cationic polyacrylamides (C-PAMs), a class of water-soluble copolymers based on acrylamide and vinyl-type comonomers with quaternary ammonium cations that are widely used in wastewater treatment and in paper industries. Linear and branched C-PAM copolymers prepared in two different polymerization methods (solution and emulsion) from varying amounts of crosslinking agent and initiator were size fractionated by FlFFF with field-programming. It was found experimentally that the linear copolymers from both polymerization methods were less than 10(8) g/mol in MW with compact, nearly spherical structures, while the branched C-PAM copolymers from the emulsion polymerization showed a significant increase in average MW up to ∼ 10(9)g/mol, which was about 20-fold greater than those from the solution method, and the branched copolymers had more compact or shrunken conformations. While both linear and branched copolymers less than 10(8) g/mol MW were well resolved in an increasing order of MW (normal mode), it was noted that branched copolymers prepared through emulsion polymerization exhibited significantly larger MWs of 10(8-)10(9) g/mol and eluted in the steric/hyperlayer mode, in which the elution order is reversed in an extreme run condition (strong initial field strength followed by a fast field decay during programming). Copyright © 2015 Elsevier B.V. All rights reserved.

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.

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.

Unsteady aerodynamics of an oscillating cascade in a compressible flow field

NASA Technical Reports Server (NTRS)

Buffum, Daniel H.; Boldman, Donald R.; Fleeter, Sanford

1987-01-01

Fundamental experiments were performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics of a cascade of biconvex airfoils executing torsion-mode oscillations at realistic reduced frequencies. Flush-mounted, high-response miniature pressure transducers were used to measure the unsteady airfoil surface pressures. The pressures were measured for three interblade phase angles at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 deg. The time-variant pressures were analyzed by means of discrete Fourier transform techniques, and these unique data were then compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle had a major effect on the chordwise distributions of the airfoil surface unsteady pressure, and that reduced frequency, incidence angle, and Mach number had a somewhat less significant effect.

The effects of opening areas on solar chimney performance

NASA Astrophysics Data System (ADS)

Ling, L. S.; Rahman, M. M.; Chu, C. M.; Misaran, M. S. bin; Tamiri, F. M.

2017-07-01

To enhance natural ventilation at day time, solar chimney is one of the suitable options for topical country like Malaysia. Solar chimney creates air flow due to stack effect caused by temperature difference between ambient and inside wall. In the solar chimney, solar energy is harvested by the inner wall that cause temperature rise compare to ambient. Therefore, the efficiency of the solar chimney depends on the availability of solar energy as well as the solar intensity. In addition, it is very hard to get good ventilation at night time by using a solar chimney. To overcome this problem one of the suitable valid option is to integrate solar chimney with turbine ventilator. A new type of solar chimney is designed and fluid flow analyzed with the computational fluid dynamics (CFD) software. The aim of CFD and theoretical study are to investigate the effect of opening areas on modified solar chimney performance. The inlet and outlet area of solar chimney are varied from 0.0224m2 to 0.6m2 and 0.1m2 to 0.14m2 respectively based on the changes of inclination angle and gap between inner and outer wall. In the CFD study the constant heat flux is considered as 500W/m2. CFD result shows that there is no significant relation between opening areas and the air flow rate through solar chimney but the ratio between inlet and outlet is significant on flow performance. If the area ratio between inlet and outlet are equal to two or larger, the performance of the solar chimney is better than the solar chimney with ratio lesser than two. The solar chimney performance does not effect if the area ratio between inlet and outlet varies from 1 to 2. This result will be useful for design and verification of actual solar chimney performance.

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.

OFF-DESIGN PERFORMANCE OF RADIAL INFLOW TURBINES

NASA Technical Reports Server (NTRS)

Wasserbauer, C. A.

1994-01-01

This program calculates off design performance of radial inflow turbines. The program uses a one dimensional solution of flow conditions through the turbine along the main streamline. The loss model accounts for stator, rotor, incidence, and exit losses. Program features include consideration of stator and rotor trailing edge blockage and computation of performance to limiting load. Stator loss (loss in kinetic energy across the stator) is proportional to the average kinetic energy in the blade row and is represented in the program by an equation which includes a stator loss coefficient determined from design point performance and then assumed to be constant for the off design calculations. Minimum incidence loss does not occur at zero incidence angle with respect to the rotor blade, but at some optimum flow angle. At high pressure ratios the level of rotor inlet velocity seemed to have an excessive influence on the loss. Using the component of velocity in the direction of the optimum flow angle gave better correlations with experimental results. Overall turbine geometry and design point values of efficiency, pressure ratio, and mass flow are needed as input information. The output includes performance and velocity diagram parameters for any number of given speeds over a range of turbine pressure ratio. The program has been implemented on the IBM 7094 and operates in batch mode.

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.

Parametric modeling and stagger angle optimization of an axial flow fan

NASA Astrophysics Data System (ADS)

Li, M. X.; Zhang, C. H.; Liu, Y.; Y Zheng, S.

2013-12-01

Axial flow fans are widely used in every field of social production. Improving their efficiency is a sustained and urgent demand of domestic industry. The optimization of stagger angle is an important method to improve fan performance. Parametric modeling and calculation process automation are realized in this paper to improve optimization efficiency. Geometric modeling and mesh division are parameterized based on GAMBIT. Parameter setting and flow field calculation are completed in the batch mode of FLUENT. A control program is developed in Visual C++ to dominate the data exchange of mentioned software. It also extracts calculation results for optimization algorithm module (provided by Matlab) to generate directive optimization control parameters, which as feedback are transferred upwards to modeling module. The center line of the blade airfoil, based on CLARK y profile, is constructed by non-constant circulation and triangle discharge method. Stagger angles of six airfoil sections are optimized, to reduce the influence of inlet shock loss as well as gas leak in blade tip clearance and hub resistance at blade root. Finally an optimal solution is obtained, which meets the total pressure requirement under given conditions and improves total pressure efficiency by about 6%.

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.

Five-hole pitot probe time-mean velocity measurements in confined swirling flows

NASA Technical Reports Server (NTRS)

Yoon, H. K.; Lilley, D. G.

1983-01-01

Nonswirling and swirling nonreacting flows in an axisymmetric test section with an expansion ratio D/d = 2, which may be equipped with contraction nozzles of area ratios 2 and 4, are investigated. The effects of a number of geometric parameters on the flow-field are investigated, among them side-wall expansion angles of 90 and 45 deg, swirl vane angles of 0, 38, 45, 60, and 70 deg, and contraction nozzle locations L/D = 1 and 2 (if present). Data are acquired by means of a five-hole pitot probe enabling three time-mean velocity components in the axial, radial, and azimuthal directions to be measured. The velocities are extensively plotted and artistic impressions of recirculation zones are set forth. The presence of a swirler is found to shorten the corner recirculation zone and to generate a central recirculation zone followed by a precessing vortex core. A gradual inlet expansion has the effect of encouraging the flow to remain close to the sidewall and shortening the extent of the corner recirculation zone in all cases investigated.

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.

Impeller flow field characterization with a laser two-focus velocimeter

NASA Astrophysics Data System (ADS)

Brozowski, L. A.; Ferguson, T. V.; Rojas, L.

1993-07-01

Use of Computational Fluid Dynamics (CFD) codes, prevalent in the rocket engine turbomachinery industry, necessitates data of sufficient quality and quantity to benchmark computational codes. Existing data bases for typical rocket engine configurations, in particular impellers, are limited. In addition, traditional data acquisition methods have several limitations: typically transducer uncertainties are 0.5% of transducer full scale and traditional pressure probes are unable to provide flow characteristics in the circumferential (blade-to-blade) direction. Laser velocimetry circumvents these limitations by providing +0.5% uncertainty in flow velocity and +0.5% uncertainty in flow angle. The percent of uncertainty in flow velocity is based on the measured value, not full range capability. The laser electronics multiple partitioning capability allows data acquired between blades as the impeller rotates, to be analyzed separately, thus providing blade-to-blade flow characterization. Unlike some probes, the non-intrusive measurements made with the laser velocimeter does not disturb the flow. To this end,, and under Contract (NAS8-38864) to the National Aeronautics and Space Administration (NASA) at Marshall Space Flight Center (MSFC), an extensive test program was undertaken at Rocketdyne. Impellers from two different generic rocket engine pump configurations were examined. The impellers represent different spectrums of pump design: the Space Shuttle Main Engine (SSME) high pressure fuel turbopump (HPFTP) impeller was designed in the 1 1970's the Consortium for CFD application in Propulsion Technology Pump Stage Technology Team (Pump Consortium) optimized impeller was designed with the aid of modern computing techniques. The tester configuration for each of the impellers consisted of an axial inlet, an inducer, a diffuser, and a crossover discharge. While the tested configurations were carefully chosen to be representative of generic rocket engine pumps, several features of both testers were intentionally atypical. A crossover discharge, downstream of the impeller, rather than a volute discharge was used to minimize asymmetric flow conditions that might be reflected in the impeller discharge flow data. Impeller shroud wear ring radial clearances were purposely close to minimize leakage flow, thus increasing confidence in using the inlet data as an input to CFD programs. The empirical study extensively examined the flow fields of the two impellers via performance of laser two-focus velocimeter surveys in an axial plane upstream of the impellers and in multiple radial planes downstream of the impellers. Both studies were performed at the impeller design flow coefficients. Inlet laser surveys that provide CFD code inlet boundary conditions were performed in one axial plane, with ten radial locations surveyed. Three wall static pressures, positioned circumferentially around the impeller inlet, were used to identify asymmetrical pressure distributions in the inlet survey plane.

Parallel computation of fluid-structural interactions using high resolution upwind schemes

NASA Astrophysics Data System (ADS)

Hu, Zongjun

An efficient and accurate solver is developed to simulate the non-linear fluid-structural interactions in turbomachinery flutter flows. A new low diffusion E-CUSP scheme, Zha CUSP scheme, is developed to improve the efficiency and accuracy of the inviscid flux computation. The 3D unsteady Navier-Stokes equations with the Baldwin-Lomax turbulence model are solved using the finite volume method with the dual-time stepping scheme. The linearized equations are solved with Gauss-Seidel line iterations. The parallel computation is implemented using MPI protocol. The solver is validated with 2D cases for its turbulence modeling, parallel computation and unsteady calculation. The Zha CUSP scheme is validated with 2D cases, including a supersonic flat plate boundary layer, a transonic converging-diverging nozzle and a transonic inlet diffuser. The Zha CUSP2 scheme is tested with 3D cases, including a circular-to-rectangular nozzle, a subsonic compressor cascade and a transonic channel. The Zha CUSP schemes are proved to be accurate, robust and efficient in these tests. The steady and unsteady separation flows in a 3D stationary cascade under high incidence and three inlet Mach numbers are calculated to study the steady state separation flow patterns and their unsteady oscillation characteristics. The leading edge vortex shedding is the mechanism behind the unsteady characteristics of the high incidence separated flows. The separation flow characteristics is affected by the inlet Mach number. The blade aeroelasticity of a linear cascade with forced oscillating blades is studied using parallel computation. A simplified two-passage cascade with periodic boundary condition is first calculated under a medium frequency and a low incidence. The full scale cascade with 9 blades and two end walls is then studied more extensively under three oscillation frequencies and two incidence angles. The end wall influence and the blade stability are studied and compared under different frequencies and incidence angles. The Zha CUSP schemes are the first time to be applied in moving grid systems and 2D and 3D calculations. The implicit Gauss-Seidel iteration with dual time stepping is the first time to be used for moving grid systems. The NASA flutter cascade is the first time to be calculated in full scale.

Design and numeric evaluation of a novel axial-flow left ventricular assist device.

PubMed

Toptop, Koral; Kadipasaoglu, Kamuran A

2013-01-01

Virtual design characteristics and performance of the first Turkish axial-flow left ventricular assist device (LVAD) are presented, with emphasis on rotor geometry. The patented rotor design includes a central flow channel carved inside the main block, which carries permanent magnets. A concentric rotor-stator gap minimizes the distance between respective magnets, improving electromagnetic efficiency and creating a second blood pathway. Dual sets of three helical blades, placed on the shaft and external surface of the rotor block, ensure unidirectionality. Hemodynamic performance was tested with computational fluid dynamics (CFD); and rotor-blade geometry was optimized, to maximize overall efficiency d and minimize backflow and wall shear stresses. For a shaft radius of 4.5 mm, rotor blade height of 2.5 mm, and blade inlet and exit metal angles of 67° and 32°, pump operation at the nominal head-flow combination (5 L/min and 100.4 mm Hg) was achieved at a rotor speed of 10,313 rpm. At the nominal point, backflow as percent of total flow was 7.29 and 29.87% at rotor inlet and exit, respectively; overall hydraulic efficiency reached 21.59%; and maximum area-averaged shroud shear was 520 Pa. Overall efficiency peaked at 24.07% for a pump flow of 6.90 L/min, and averaged at 22.57% within the flow range of 4-8 L/min. We concluded that the design satisfies initial rotor design criteria, and that continued studies with diffuser optimization and transient flow analysis are warranted.

Exit blade geometry and part-load performance of small axial flow propeller turbines: An experimental investigation

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

Singh, Punit; Nestmann, Franz

2010-09-15

A detailed experimental investigation of the effects of exit blade geometry on the part-load performance of low-head, axial flow propeller turbines is presented. Even as these turbines find important applications in small-scale energy generation using micro-hydro, the relationship between the layout of blade profile, geometry and turbine performance continues to be poorly characterized. The experimental results presented here help understand the relationship between exit tip angle, discharge through the turbine, shaft power, and efficiency. The modification was implemented on two different propeller runners and it was found that the power and efficiency gains from decreasing the exit tip angle couldmore » be explained by a theoretical model presented here based on classical theory of turbomachines. In particular, the focus is on the behaviour of internal parameters like the runner loss coefficient, relative flow angle at exit, mean axial flow velocity and net tangential flow velocity. The study concluded that the effects of exit tip modification were significant. The introspective discussion on the theoretical model's limitation and test facility suggests wider and continued experimentation pertaining to the internal parameters like inlet vortex profile and exit swirl profile. It also recommends thorough validation of the model and its improvement so that it can be made capable for accurate characterization of blade geometric effects. (author)« less

Effects of front-loading and stagger angle on endwall losses of high lift low pressure turbine vanes

NASA Astrophysics Data System (ADS)

Lyall, M. Eric

Past efforts to reduce the airfoil count in low pressure turbines have produced high lift profiles with unacceptably high endwall loss. The purpose of the current work is to suggest alternative approaches for reducing endwall losses. The effects of the fluid mechanics and high lift profile geometry are considered. Mixing effects of the mean flow and turbulence fields are decoupled to show that mean flow shear in the endwall wake is negligible compared to turbulent shear, indicating that turbulence dissipation is the primary cause of total pressure loss. The mean endwall flow field does influence total pressure loss by causing excessive wake growth and perhaps outright separation on the suction surface. For equivalent stagger angles, a front-loaded high lift profile will produce less endwall loss than one aft-loaded, primarily by suppressing suction surface flow separation. Increasing the stagger setting, however, increases the endwall loss due to the static pressure field generating a stronger blockage relative to the incoming endwall boundary layer flow and causing a larger mass of fluid to become entrained in the horseshoe vortex. In short, front-loading the pressure distribution suppresses suction surface separation whereas limiting the stagger angle suppresses inlet boundary layer separation. Results of this work suggest that a front-loaded low stagger profile be used at the endwall to reduce the endwall loss.

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.

Analysis of Fuel Vaporization, Fuel-Air Mixing, and Combustion in Integrated Mixer-Flame Holders

NASA Technical Reports Server (NTRS)

Deur, J. M.; Cline, M. C.

2004-01-01

Requirements to limit pollutant emissions from the gas turbine engines for the future High-Speed Civil Transport (HSCT) have led to consideration of various low-emission combustor concepts. One such concept is the Integrated Mixer-Flame Holder (IMFH). This report describes a series of IMFH analyses performed with KIVA-II, a multi-dimensional CFD code for problems involving sprays, turbulence, and combustion. To meet the needs of this study, KIVA-II's boundary condition and chemistry treatments are modified. The study itself examines the relationships between fuel vaporization, fuel-air mixing, and combustion. Parameters being considered include: mixer tube diameter, mixer tube length, mixer tube geometry (converging-diverging versus straight walls), air inlet velocity, air inlet swirl angle, secondary air injection (dilution holes), fuel injection velocity, fuel injection angle, number of fuel injection ports, fuel spray cone angle, and fuel droplet size. Cases are run with and without combustion to examine the variations in fuel-air mixing and potential for flashback due to the above parameters. The degree of fuel-air mixing is judged by comparing average, minimum, and maximum fuel/air ratios at the exit of the mixer tube, while flame stability is monitored by following the location of the flame front as the solution progresses from ignition to steady state. Results indicate that fuel-air mixing can be enhanced by a variety of means, the best being a combination of air inlet swirl and a converging-diverging mixer tube geometry. With the IMFH configuration utilized in the present study, flashback becomes more common as the mixer tube diameter is increased and is instigated by disturbances associated with the dilution hole flow.

Computational Analyses of the LIMX TBCC Inlet High-Speed Flowpath

NASA Technical Reports Server (NTRS)

Dippold, Vance F., III

2012-01-01

Reynolds-Averaged Navier-Stokes (RANS) simulations were performed for the high-speed flowpath and isolator of a dual-flowpath Turbine-Based Combined-Cycle (TBCC) inlet using the Wind-US code. The RANS simulations were performed in preparation for the Large-scale Inlet for Mode Transition (LIMX) model tests in the NASA Glenn Research Center (GRC) 10- by 10-ft Supersonic Wind Tunnel. The LIMX inlet has a low-speed flowpath that is coupled to a turbine engine and a high-speed flowpath designed to be coupled to a Dual-Mode Scramjet (DMSJ) combustor. These RANS simulations were conducted at a simulated freestream Mach number of 4.0, which is the nominal Mach number for the planned wind tunnel testing with the LIMX model. For the simulation results presented in this paper, the back pressure, cowl angles, and freestream Mach number were each varied to assess the performance and robustness of the high-speed inlet and isolator. Under simulated wind tunnel conditions at maximum inlet mass flow rates, the high-speed flowpath pressure rise was found to be greater than a factor of four. Furthermore, at a simulated freestream Mach number of 4.0, the high-speed flowpath and isolator showed stability for freestream Mach number that drops 0.1 Mach below the design point. The RANS simulations indicate the yet-untested highspeed inlet and isolator flowpath should operate as designed. The RANS simulation results also provided important insight to researchers as they developed test plans for the LIMX experiment in GRC s 10- by 10-ft Supersonic Wind Tunnel.

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.

A computer program for the calculation of the flow field in supersonic mixed-compression inlets at angle of attack using the three-dimensional method of characteristics with discrete shock wave fitting

NASA Technical Reports Server (NTRS)

Vadyak, J.; Hoffman, J. D.; Bishop, A. R.

1978-01-01

The calculation procedure is based on the method of characteristics for steady three-dimensional flow. The bow shock wave and the internal shock wave system were computed using a discrete shock wave fitting procedure. The general structure of the computer program is discussed, and a brief description of each subroutine is given. All program input parameters are defined, and a brief discussion on interpretation of the output is provided. A number of sample cases, complete with data deck listings, are presented.

Artificial neural networks modelling the prednisolone nanoprecipitation in microfluidic reactors.

PubMed

Ali, Hany S M; Blagden, Nicholas; York, Peter; Amani, Amir; Brook, Toni

2009-06-28

This study employs artificial neural networks (ANNs) to create a model to identify relationships between variables affecting drug nanoprecipitation using microfluidic reactors. The input variables examined were saturation levels of prednisolone, solvent and antisolvent flow rates, microreactor inlet angles and internal diameters, while particle size was the single output. ANNs software was used to analyse a set of data obtained by random selection of the variables. The developed model was then assessed using a separate set of validation data and provided good agreement with the observed results. The antisolvent flow rate was found to have the dominant role on determining final particle size.

Effect of simulated forward airspeed on small-scale-model externally blown flap noise

NASA Technical Reports Server (NTRS)

Goodykoontz, J. H.; Dorsch, R. G.; Olsen, W. A.

1976-01-01

Noise tests were conducted on a small-scale model of an externally blown flap lift augmentation system. The nozzle/wing model was subjected to external flow that simulated takeoff and landing flight velocities by placing it in a 33-centimeter-diameter free jet. The results showed that external flow attenuated the noise associated with the various configurations tested. The amount of attenuation depended on flap setting. More attenuation occurred with a trailing-flap setting of 20 deg than with one of 60 deg. Noise varied with relative velocity as a function of the trailing-flap setting and the angle from the nozzle inlet.

Research on Supersonic Inlet Bleed

NASA Technical Reports Server (NTRS)

Davis, David O.; Vyas, Manan A.; Slater, John W.

2012-01-01

Phase I data results of the Fundamental Inlet Bleed Experiments project at NASA Glenn Research Center (GRC) are presented which include flow coefficient results for two single-hole boundary-layer bleed configurations. The bleed configurations tested are round holes at inclination angles of 90deg and 20deg both having length-to-diameter ratios of 2.0. Results were obtained at freestream Mach numbers of 1.33, 1.62, 1.98, 2.46, and 2.92 and unit Reynolds numbers of 0.984, 1.89, and 2.46 10(exp 7)/m. Approach boundary-layer data are presented for each flow condition and the flow coefficient results are compared to existing multi-hole data obtained under similar conditions. For the 90deg hole, the single and multi-hole distributions agree fairly well with the exception that under supercritical operation, the multi-hole data chokes at higher flow coefficient levels. This behavior is also observed for the 20deg hole but to a lesser extent. The 20deg hole also shows a markedly different characteristic at subcritical operation. Also presented are preliminary results of a Computational Fluid Dynamics (CFD) analysis of both configurations at the Mach 1.33 and a unit Reynolds number of 2.46 10(exp 7)/m. Comparison of the results shows the agreement to be very good.

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.

Experimental Investigation of Actuators for Flow Control in Inlet Ducts

NASA Astrophysics Data System (ADS)

Vaccaro, John; Elimelech, Yossef; Amitay, Michael

2010-11-01

Attractive to aircraft designers are compact inlets, which implement curved flow paths to the compressor face. These curved flow paths could be employed for multiple reasons. One of which is to connect the air intake to the engine embedded in the aircraft body. A compromise must be made between the compactness of the inlet and its aerodynamic performance. The aerodynamic purpose of inlets is to decelerate the oncoming flow before reaching the engine while minimizing total pressure loss, unsteadiness and distortion. Low length-to-diameter ratio inlets have a high degree of curvature, which inevitably causes flow separation and secondary flows. Currently, the length of the propulsion system is constraining the overall size of Unmanned Air Vehicles (UAVs), thus, smaller more efficient aircrafts could be realized if the propulsion system could be shortened. Therefore, active flow control is studied in a compact (L/D=1.5) inlet to improve performance metrics. Actuation from a spanwise varying coanda type ejector actuator and a hybrid coanda type ejector / vortex generator jet actuator is investigated. Special attention will be given to the pressure recovery at the AIP along with unsteady pressure signatures along the inlet surface and at the AIP.

Flow visualization studies of VTOL aircraft models during Hover in ground effect

NASA Technical Reports Server (NTRS)

Mourtos, Nikos J.; Couillaud, Stephane; Carter, Dale; Hange, Craig; Wardwell, Doug; Margason, Richard J.

1995-01-01

A flow visualization study of several configurations of a jet-powered vertical takeoff and landing (VTOL) aircraft model during hover in ground effect was conducted. A surface oil flow technique was used to observe the flow patterns on the lower surfaces of the model. There were significant configuration effects. Wing height with respect to fuselage, the presence of an engine inlet duct beside the fuselage, and nozzle pressure ratio are seen to have strong effects on the surface flow angles on the lower surface of the wing. This test was part of a program to improve the methods for predicting the hot gas ingestion (HGI) for jet-powered vertical/short takeoff and landing (V/STOL) aircraft. The tests were performed at the Jet Calibration and Hover Test (JCAHT) Facility at Ames Research Center.

Low speed test of a high-bypass-ratio propulsion system with an asymmetric inlet designed for a tilt-nacelle V/STOL airplane

NASA Technical Reports Server (NTRS)

Syberg, J.

1978-01-01

A large scale model of a lift/cruise fan inlet designed for a tilt nacelle V/STOL airplane was tested with a high bypass ratio turbofan. Testing was conducted at low freestream velocities with inlet angles of attack ranging from 0 deg to 120 deg. The operating limits for the nacelle were found to be related to inlet boundary layer separation. Small separations originating in the inlet diffuser cause little or no performance degradation. However, at sufficiently severe freestream conditions the separation changes abruptly to a lip separation. This change is associated with a significant reduction in nacelle net thrust as well as a sharp increase in fan blade vibratory stresses. Consequently, the onset of lip separation is regarded as the nacelle operating limit. The test verified that the asymmetric inlet design will provide high performance and stable operation at the design forward speed and angle of attack conditions. At some of these, however, operation near the lower end of the design inlet airflow range is not feasible due to the occurrence of lip separation.

Investigation of Inlet Control Parameters for an External-internal-compression Inlet from Mach 2.1 to 3.0

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Bowditch, D. N.

1958-01-01

Investigation of the control parameters of an external-internal compression inlet indicates that the cowl-lip shock provides a signal to position the spike and to start the inlet over a Mach number range from 2.1 to 3.0. Use of a single fixed probe position to control the spike over the range of conditions resulted in a 3.7-count loss in total-pressure recovery at Mach 3.0 and 0 deg angle of attack. Three separate shock-sensing-probe positions were required to set the spike for peak recovery from Mach 2.1 to 3.0 and angles of attack from 0 deg to 6 deg. When the inlet was unstarted, an erroneous signal was obtained from the normal-shock control through most of the starting cycle that prevented the inlet from starting. Therefore, it was necessary to over-ride the normal-shock control signal and not allow the control to position the terminal shock until the spike was positioned.

An inverse inviscid method for the design of quasi-three dimensional rotating turbomachinery cascades

NASA Technical Reports Server (NTRS)

Bonataki, E.; Chaviaropoulos, P.; Papailiou, K. D.

1991-01-01

A new inverse inviscid method suitable for the design of rotating blade sections lying on an arbitrary axisymmetric stream-surface with varying streamtube width is presented. The geometry of the axisymmetric stream-surface and the streamtube width variation with meridional distance, the number of blades, the inlet flow conditions, the rotational speed and the suction and pressure side velocity distributions as functions of the normalized arc-length are given. The flow is considered irrotational in the absolute frame of reference and compressible. The output of the computation is the blade section that satisfies the above data. The method solves the flow equations on a (phi 1, psi) potential function-streamfunction plane for the velocity modulus, W and the flow angle beta; the blade section shape can then be obtained as part of the physical plane geometry by integrating the flow angle distribution along streamlines. The (phi 1, psi) plane is defined so that the monotonic behavior of the potential function is guaranteed, even in cases with high peripheral velocities. The method is validated on a rotating turbine case and used to design new blades. To obtain a closed blade, a set of closure conditions were developed and referred.

Experimental study of low aspect ratio compressor blading

NASA Technical Reports Server (NTRS)

Reid, L.; Moore, R. D.

1979-01-01

The effects of low aspect ratio blading on aerodynamic performance were examined. Four individual transonic compressor stages, representative of the inlet stage of an advanced high pressure ratio core compressor, are discussed. The flow phenomena for the four stages are investigated. Comparisons of blade element parameters are presented for the two different aspect ratio configurations. Blade loading levels are compared for the near stall conditions and comparisons are made of loss and diffusion factors over the operating range of incidence angles.

Gust Response Analysis of a Turbine Cascade

NASA Technical Reports Server (NTRS)

Gorla, R. S. R.; Reddy, T. S. R.; Reddy, D. R.; Kurkov, A. P.

2001-01-01

A study was made of the gust response of an annular turbine cascade using a two-dimensional Navier Stokes code. The time-marching CFD code, NPARC, was used to calculate the unsteady forces due to the fluid flow. The computational results were compared with a previously published experimental data for the annular cascade reported in the literature. Reduced frequency, Mach number and angle of incidence were varied independently and the gust velocity was sinusoidal. For the high inlet velocity case, the cascade was nearly choked.

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.

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.

Numerical simulation of the flow about the F-18 HARV at high angle of attack

NASA Technical Reports Server (NTRS)

Murman, Scott M.

1994-01-01

This report summarizes research done over the past two years as part of NASA Grant NCC 2-729. This research has been aimed at validating numerical methods for computing the flow about the complete F-18 HARV at alpha = 30 deg and alpha = 45 deg. At 30 deg angle of attack, the flow about the F-18 is dominated by the formation, and subsequent breakdown, of strong vortices over the wing leading-edge extensions (LEX). As the angle of attack is increased to alpha = 45 deg, the fuselage forebody of the F-18 contains significant laminar and transitional regions which are not present at alpha = 30 deg. Further, the flow over the LEX at alpha = 45 deg is dominated by an unsteady shedding in time, rather than strong coherent vortices. This complex physics, combined with the complex geometry of a full aircraft configuration, provides a challenge for current computational fluid dynamics (CFD) techniques. The following sections present the numerical method and grid generation scheme that was used, a review of prior research done to numerically model the F-18 HARV, and a discussion of the current research. The current research is broken into two main topics: the effect of engine-inlet mass-flow rate on the F-18 vortex breakdown position, and the results using a refined F-18 computational model to compute the flow at alpha = 30 deg and alpha = 45 deg.

Numerical simulation of the flow about the F-18 HARV at high angle of attack

NASA Technical Reports Server (NTRS)

Murman, Scott M.

1995-01-01

This research has been aimed at validating numerical methods for computing the flow about the complete F-18 HARV at alpha = 30 deg and alpha = 45 deg. At 30 deg angle of attack, the flow about the F-18 is dominated by the formation, and subsequent breakdown, of strong vortices over the wing leading-edge extensions (LEX). As the angle of attack is increased to alpha = 45 deg, the fuselage forebody of the F-18 contains significant laminar and transitional regions which are not present at alpha = 30 deg. Further, the flow over the LEX at alpha = 45 deg is dominated by an unsteady shedding in time, rather than strong coherent vortices. This complex physics, combined with the complex geometry of a full-aircraft configuration, provides a challenge for current computational fluid dynamics (CFD) techniques. The following sections present the numerical method and grid generation scheme that was used, a review of prior research done to numerically model the F-18 HARV, and a discussion of the current research. The current research is broken into three main topics; the effect of engine-inlet mass-flow rate on the F-18 vortex breakdown position, the results using a refined F-18 computational model to compute the flow at alpha = 30 deg and alpha = 45 deg, and research done using the simplified geometry of an ogive-cylinder configuration to investigate the physics of unsteady shear-layer shedding. The last section briefly summarizes the discussion.

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.

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.

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.

Sample of CFD optimization of a centrifugal compressor stage

NASA Astrophysics Data System (ADS)

Galerkin, Y.; Drozdov, A.

2015-08-01

Industrial centrifugal compressor stage is a complicated object for gas dynamic design when the goal is to achieve maximum efficiency. The Authors analyzed results of CFD performance modeling (NUMECA Fine Turbo calculations). Performance prediction in a whole was modest or poor in all known cases. Maximum efficiency prediction was quite satisfactory to the contrary. Flow structure in stator elements was in a good agreement with known data. The intermediate type stage “3D impeller + vaneless diffuser+ return channel” was designed with principles well proven for stages with 2D impellers. CFD calculations of vaneless diffuser candidates demonstrated flow separation in VLD with constant width. The candidate with symmetrically tampered inlet part b3 / b2 = 0,73 appeared to be the best. Flow separation takes place in the crossover with standard configuration. The alternative variant was developed and numerically tested. The obtained experience was formulated as corrected design recommendations. Several candidates of the impeller were compared by maximum efficiency of the stage. The variant with gas dynamic standard principles of blade cascade design appeared to be the best. Quasi - 3D non-viscid calculations were applied to optimize blade velocity diagrams - non-incidence inlet, control of the diffusion factor and of average blade load. “Geometric” principle of blade formation with linear change of blade angles along its length appeared to be less effective. Candidates’ with different geometry parameters were designed by 6th math model version and compared. The candidate with optimal parameters - number of blades, inlet diameter and leading edge meridian position - is 1% more effective than the stage of the initial design.

Are two different projections of the inlet view necessary for the percutaneous placement of iliosacral screws?

PubMed

Ozmeric, A; Yucens, M; Gultaç, E; Açar, H I; Aydogan, N H; Gül, D; Alemdaroglu, K B

2015-05-01

We hypothesised that the anterior and posterior walls of the body of the first sacral vertebra could be visualised with two different angles of inlet view, owing to the conical shape of the sacrum. Six dry male cadavers with complete pelvic rings and eight dry sacrums with K-wires were used to study the effect of canting (angling the C-arm) the fluoroscope towards the head in 5° increments from 10° to 55°. Fluoroscopic images were taken in each position. Anterior and posterior angles of inclination were measured between the upper sacrum and the vertical line on the lateral view. Three authors separately selected the clearest image for overlapping anterior cortices and the upper sacral canal in the cadaveric models. The dry bone and K-wire models were scored by the authors, being sure to check whether the K-wire was in or out. In the dry bone models the mean score of the relevant inlet position of the anterior or posterior inclination was 8.875 (standard deviation (sd) 0.35), compared with the inlet position of the opposite inclination of -5.75 (sd 4.59). We found that two different inlet views should be used separately to evaluate the borders of the body of the sacrum using anterior and posterior inclination angles of the sacrum, during placement of iliosacral screws. ©2015 The British Editorial Society of Bone & Joint Surgery.

Separation control in adverse pressure gradients using high-speed microjets

NASA Astrophysics Data System (ADS)

Kumar, Vikas

Inlets to aircraft propulsion systems must supply flow to the compressor with minimal pressure loss, flow distortion or unsteadiness. Flow separation in internal flows such as inlets and ducts in aircraft propulsion systems and external flows such as over aircraft wings, is undesirable as it reduces the overall system performance. The objective of present study is to understand the nature of separation and more importantly, to explore the applicability of high-speed microjets to actively control this flow separation. The geometry used for this experimental study was a generic backward facing "Stratford Ramp" equipped with arrays of high-speed microjets. The incoming flow was examined over a freestream velocity range of 10-65m/s and at ramp angle in range of 0-10°. It was observed that the flow separates at 30m/s and beyond for all angle of attack. The magnitude and extent of separation bubble increases with increasing adverse pressure gradients and/or increase in free-stream velocity. The separated flow for all the examined conditions was completely attached using suitable array of high-speed microjets. The most notable fact was that elimination of reverse velocity regions was accompanied by a reduction in flow unsteadiness and increased two-dimensionality in the flow. In particular, these gains were achieved with a minimal mass flux, less than 0.2% of the primary flow based on 30% Boundary Layer Ingesting duct. Detailed measurements were obtained to understand the flow control dynamics. The control effectiveness was found to be dependent on the actuation location with respect to separation, jet to cross-flow momentum ratio and the angle at which microjets supply the momentum. It was also determined that the control effect of the microjets, in part, is due to creation of strong stream-wise vortices which enhance the mixing between low-momentum fluid closer to the surface and high-momentum fluid further away from the surface. The penetration depth of microjets was found to be much higher than that of a jet exiting in to uniform cross-flow and correlations were developed to predict this. Subsequently, means for identification of the flow conditions were sought to develop a simple, robust, complete control strategy. It was observed that the flow conditions were very well represented in unsteady surface pressure measurements. The unsteady surface pressure and velocity field were correlated to develop a simple scheme to predict the peak unsteadiness location over the surface. The results from this model and knowledge of microjet in cross flow was used to provide guidelines for an active control strategy. A case study was then undertaken to validate the results obtained using the model. The results show that the model is a good first step towards developing a simple, robust, active-adaptive separation control strategy using microjets.

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.

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.

Local Flow Conditions for Propulsion Experiments on the NASA F-15B Propulsion Flight Test Fixture

NASA Technical Reports Server (NTRS)

Vachon, Michael J.; Moes, Timothy R.; Corda, Stephen

2005-01-01

Local flow conditions were measured underneath the National Aeronautics and Space Administration F-15B airplane to support development of future experiments on the Propulsion Flight Test Fixture (PFTF). The local Mach number and flow angles were measured using a conventional air data boom on a cone-cylinder mounted under the PFTF and compared with the airplane air data nose boom measurements. At subsonic flight speeds, the airplane and PFTF Mach numbers were approximately equal. Transonic Mach number values were up to 0.1 greater at the PFTF than the airplane, which is a counterintuitive result. The PFTF local supersonic Mach numbers were as much as 0.46 less than the airplane values. The maximum local Mach number at the PFTF was approximately 1.6 at an airplane Mach number near 2.0. The PFTF local angle of attack was negative at all Mach numbers, ranging from -3 to -8 degrees. When the airplane angle of sideslip was zero, the PFTF local value was zero between Mach 0.8 and Mach 1.1, -2 degrees between Mach 1.1 and Mach 1.5, and increased from zero to 1 degree from Mach 1.5 to Mach 2.0. Airplane inlet shock waves crossed the aerodynamic interface plane between Mach 1.85 and Mach 1.90.

Fundamental Study of a Single Point Lean Direct Injector. Part I: Effect of Air Swirler Angle and Injector Tip Location on Spray Characteristics

NASA Technical Reports Server (NTRS)

Tedder, Sarah A.; Hicks, Yolanda R.; Tacina, Kathleen M.; Anderson, Robert C.

2014-01-01

Lean direct injection (LDI) is a combustion concept to reduce oxides of nitrogen (NOx) for next generation aircraft gas turbine engines. These newer engines have cycles that increase fuel efficiency through increased operating pressures, which increase combustor inlet temperatures. NOx formation rates increase with higher temperatures; the LDI strategy avoids high temperature by staying fuel lean and away from stoichiometric burning. Thus, LDI relies on rapid and uniform fuel/air mixing. To understand this mixing process, a series of fundamental experiments are underway in the Combustion and Dynamics Facility at NASA Glenn Research Center. This first set of experiments examines cold flow (non-combusting) mixing using air and water. Using laser diagnostics, the effects of air swirler angle and injector tip location on the spray distribution, recirculation zone, and droplet size distribution are examined. Of the three swirler angles examined, 60 deg is determined to have the most even spray distribution. The injector tip location primarily shifts the flow without changing the structure, unless the flow includes a recirculation zone. When a recirculation zone is present, minimum axial velocity decreases as the injector tip moves downstream towards the venturi exit; also the droplets become more uniform in size and angular distribution.

Fundamental Study of a Single Point Lean Direct Injector. Part I: Effect of Air Swirler Angle and Injector Tip Location on Spray Characteristics

NASA Technical Reports Server (NTRS)

Tedder, Sarah A.; Hicks, Yolanda R.; Tacina, Kathleen M.; Anderson, Robert C.

2015-01-01

Lean direct injection (LDI) is a combustion concept to reduce oxides of nitrogen (NOx) for next generation aircraft gas turbine engines. These newer engines have cycles that increase fuel efficiency through increased operating pressures, which increase combustor inlet temperatures. NOx formation rates increase with higher temperatures; the LDI strategy avoids high temperature by staying fuel lean and away from stoichiometric burning. Thus, LDI relies on rapid and uniform fuel/air mixing. To understand this mixing process, a series of fundamental experiments are underway in the Combustion and Dynamics Facility at NASA Glenn Research Center. This first set of experiments examines cold flow (non-combusting) mixing using air and water. Using laser diagnostics, the effects of air swirler angle and injector tip location on the spray distribution, recirculation zone, and droplet size distribution are examined. Of the three swirler angles examined, 60 degrees is determined to have the most even spray distribution. The injector tip location primarily shifts the flow without changing the structure, unless the flow includes a recirculation zone. When a recirculation zone is present, minimum axial velocity decreases as the injector tip moves downstream towards the venturi exit; also the droplets become more uniform in size and angular distribution.

Heat transfer and flow characteristics on a gas turbine shroud.

PubMed

Obata, M; Kumada, M; Ijichi, N

2001-05-01

The work described in this paper is an experimental investigation of the heat transfer from the main flow to a turbine shroud surface, which may be applicable to ceramic gas turbines. Three kinds of turbine shrouds are considered with a flat surface, a taper surface and a spiral groove surface opposite to the blades in an axial flow turbine of actual turbo-charger. Heat transfer measurements were performed for the experimental conditions of a uniform heat flux or a uniform wall temperature. The effects of the inlet flow angle, rotational speed, and tip clearance on the heat transfer coefficient were clarified under on- and off-design flow conditions. The mean heat transfer coefficient was correlated to the blade Reynolds number and tip clearance, and compared with an experimental correlation and measurements of a flat surface. A comparison was also made for the measurement of static pressure distributions.

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.

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.

Effect of underwing aft-mounted nacelles on the longitudinal aerodynamic characteristics of a high-wing transport airplane

NASA Technical Reports Server (NTRS)

Abeyounis, W. K.; Patterson, J. C., Jr.

1985-01-01

As part of a propulsion/airframe integration program, tests were conducted in the Langley 16-Foot Transonic Tunnel to determine the longitudinal aerodynamic effects of installing flow through engine nacelles in the aft underwing position of a high wing transonic transfer airplane. Mixed flow nacelles with circular and D-shaped inlets were tested at free stream Mach numbers from 0.70 to 0.85 and angles of attack from -2.5 deg to 4.0 deg. The aerodynamic effects of installing antishock bodies on the wing and nacelle upper surfaces as a means of attaching and supporting nacelles in an extreme aft position were investigated.

Flow Environment Study Near the Empennage of a 15-Percent Scale Helicopter Model

NASA Technical Reports Server (NTRS)

Gorton, Susan Althoff; Berry, John D.; Hodges, W. Todd; Reis, Deane G.

2000-01-01

Development of advanced rotorcraft configurations has highlighted a need for high-quality experimental data to support the development of flexible and accurate analytical design tools. To provide this type of data, a test program was conducted in the Langley 14- by 22-Foot Subsonic Tunnel to measure the flow near the empennage of a 15-percent scale powered helicopter model with an operating tail fan. Three-component velocity profiles were measured with laser velocimetry (LV) one chord forward of the horizontal tail for four advance ratios to evaluate the effect of the rotor wake impingement on the horizontal tail angle of attack. These velocity data indicate the horizontal tail can experience unsteady angle of attack variations of over 30 degrees due to the rotor wake influence. The horizontal tail is most affected by the rotor wake above advance ratios of 0.10. Velocity measurements of the flow on the inlet side of the tail fan were made for a low-speed flight condition using conventional LV techniques. The velocity data show an accelerated flow near the tail fan duct, and vorticity calculations track the passage of main rotor wake vortices through the measurement plane.

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.

Tactical missile aerodynamics

NASA Technical Reports Server (NTRS)

Hemsch, Michael J. (Editor); Nielsen, Jack N. (Editor)

1986-01-01

The present conference on tactical missile aerodynamics discusses autopilot-related aerodynamic design considerations, flow visualization methods' role in the study of high angle-of-attack aerodynamics, low aspect ratio wing behavior at high angle-of-attack, supersonic airbreathing propulsion system inlet design, missile bodies with noncircular cross section and bank-to-turn maneuvering capabilities, 'waverider' supersonic cruise missile concepts and design methods, asymmetric vortex sheding phenomena from bodies-of-revolution, and swept shock wave/boundary layer interaction phenomena. Also discussed are the assessment of aerodynamic drag in tactical missiles, the analysis of supersonic missile aerodynamic heating, the 'equivalent angle-of-attack' concept for engineering analysis, the vortex cloud model for body vortex shedding and tracking, paneling methods with vorticity effects and corrections for nonlinear compressibility, the application of supersonic full potential method to missile bodies, Euler space marching methods for missiles, three-dimensional missile boundary layers, and an analysis of exhaust plumes and their interaction with missile airframes.

Flow-structure interaction simulation of voice production in a canine larynx

NASA Astrophysics Data System (ADS)

Jiang, Weili; Zheng, Xudong; Xue, Qian; Oren, Liran; Khosla, Sid

2017-11-01

Experimental measurements conducted on a hemi-larynx canine vocal fold showed that negative pressures formed in the glottis near the superior surface of the vocal fold in the closing phase even without a supra-glottal vocal tract. It was hypothesized that such negative pressures were due to intraglottal vortices caused by flow separation in a divergent vocal tract during vocal fold closing phase. This work aims to test this hypothesis from the numerical aspect. Flow-structure interaction simulations are performed in realistic canine laryngeal shapes. In the simulations, a sharp interface immersed boundary method based incompressible flow solver is utilized to model the air flow; a finite element based solid mechanics solver is utilized to model the vocal fold vibration. The geometric structure of the vocal fold and vocal tract are based on MRI scans of a mongrel canine. The vocal fold tissue is modeled as transversely isotropic nonlinear materials with a vertical stiffness gradient. Numerical indentation is first performed and compared with the experiment data to obtain the material properties. Simulation setup about the inlet and outlet pressure follows the setup in the experiment. Simulation results including the fundamental frequency, air flow rate, the divergent angle will be compared with the experimental data, providing the validation of the simulation approach. The relationship between flow separation, intra-glottal vortices, divergent angle and flow rate will be comprehensively analyzed.

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.

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

Selected results of the F-15 propulsion interactions program

NASA Technical Reports Server (NTRS)

Webb, L. D.; Nugent, J.

1982-01-01

A better understanding of propulsion system/airframe flow interactions could aid in the reduction of aircraft drag. For this purpose, NASA and the United States Air Force have conducted a series of wind-tunnel and flight tests on the F-15 airplane. This paper presents a correlation of flight test data from tests conducted at the NASA Dryden Flight Research Facility of the Ames Research Center, with data obtained from wind-tunnel tests. Flights were made at stabilized Mach numbers around 0.6, 0.9, 1.2, and 1.5 with accelerations up to near Mach number 2. Wind-tunnel tests used a 7.5 percent-scale F-15 inlet/airframe model. Flight and wind-tunnel pressure coefficients showed good agreement in most cases. Correlation of interaction effects caused by changes in cowl angle, angle-of-attack, and Mach number are presented. For the afterbody region, the pressure coefficients on the nozzle surfaces were influenced by boattail angles and Mach number. Boundary-layer thickness decreased as angle of attack increased above 4 deg.

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.

Far-field noise and internal modes from a ducted propeller at simulated aircraft takeoff conditions

NASA Astrophysics Data System (ADS)

Woodward, Richard P.; Bock, Lawrence A.; Heidelberg, Laurence J.; Hall, David G.

1992-01-01

The ducted propeller offers structural and acoustic benefits typical of conventional turbofan engines while retaining much of the aeroacoustic benefits of the unducted propeller. A model Advanced Ducted Propeller (ADP) was tested in the NASA Lewis Low-Speed Anechoic Wind Tunnel at a simulated takeoff velocity of Mach 0.2. The ADP model was designed and manufactured by the Pratt and Whitney Division of United Technologies. The 16-blade rotor ADP was tested with 22- and 40-vane stators to achieve cut-on and cut-off criterion with respect to propagation of the fundamental rotor-stator interaction tone. Additional test parameters included three inlet lengths, three nozzle sizes, two spinner configurations, and two rotor rub strip configurations. The model was tested over a range of rotor blade setting angles and propeller axis angles-of-attack. Acoustic data were taken with a sideline translating microphone probe and with a unique inlet microphone probe which identified inlet rotating acoustic modes. The beneficial acoustic effects of cut-off were clearly demonstrated. A 5 dB fundamental tone reduction was associated with the long inlet and 40-vane sector, which may relate to inlet propeller axis angle-of-attack at rotor speeds of at least 96 percent design.

Far-field noise and internal modes from a ducted propeller at simulated aircraft takeoff conditions

NASA Astrophysics Data System (ADS)

Woodward, Richard P.; Bock, Lawrence A.; Heidelberg, Laurence J.; Hall, David G.

The ducted propeller offers structural and acoustic benefits typical of conventional turbofan engines while retaining much of the aeroacoustic benefits of the unducted propeller. A model Advanced Ducted Propeller (ADP) was tested in the NASA Lewis Low-Speed Anechoic Wind Tunnel at a simulated takeoff velocity of Mach 0.2. The ADP model was designed and manufactured by the Pratt and Whitney Division of United Technologies. The 16-blade rotor ADP was tested with 22- and 40-vane stators to achieve cut-on and cut-off criterion with respect to propagation of the fundamental rotor-stator interaction tone. Additional test parameters included three inlet lengths, three nozzle sizes, two spinner configurations, and two rotor rub strip configurations. The model was tested over a range of rotor blade setting angles and propeller axis angles-of-attack. Acoustic data were taken with a sideline translating microphone probe and with a unique inlet microphone probe which identified inlet rotating acoustic modes. The beneficial acoustic effects of cut-off were clearly demonstrated. A 5 dB fundamental tone reduction was associated with the long inlet and 40-vane sector, which may relate to inlet duct geometry. The fundamental tone level was essentially unaffected by propeller axis angle-of-attack at rotor speeds of at least 96 percent design.

Far-field noise and internal modes from a ducted propeller at simulated aircraft takeoff conditions

NASA Technical Reports Server (NTRS)

Woodward, Richard P.; Bock, Lawrence A.; Heidelberg, Laurence J.; Hall, David G.

1992-01-01

The ducted propeller offers structural and acoustic benefits typical of conventional turbofan engines while retaining much of the aeroacoustic benefits of the unducted propeller. A model Advanced Ducted Propeller (ADP) was tested in the NASA Lewis Low-Speed Anechoic Wind Tunnel at a simulated takeoff velocity of Mach 0.2. The ADP model was designed and manufactured by the Pratt and Whitney Division of United Technologies. The 16-blade rotor ADP was tested with 22- and 40-vane stators to achieve cut-on and cut-off criterion with respect to propagation of the fundamental rotor-stator interaction tone. Additional test parameters included three inlet lengths, three nozzle sizes, two spinner configurations, and two rotor rub strip configurations. The model was tested over a range of rotor blade setting angles and propeller axis angles-of-attack. Acoustic data were taken with a sideline translating microphone probe and with a unique inlet microphone probe which identified inlet rotating acoustic modes. The beneficial acoustic effects of cut-off were clearly demonstrated. A 5 dB fundamental tone reduction was associated with the long inlet and 40-vane sector, which may relate to inlet duct geometry. The fundamental tone level was essentially unaffected by propeller axis angle-of-attack at rotor speeds of at least 96 percent design.

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.

Investigation of an underslung half-cone inlet with compression surface mounted outboard from fuselage at Mach numbers of 1.5, 1.8, and 2.0

NASA Technical Reports Server (NTRS)

Yeager, Richard A; Gertsma, Laurence W

1958-01-01

An investigation was conducted to determine the performance of an underslung half-cone inlet mounted on a missile forebody model with the compression surface outboard from the fuselage. The inlet was designed for shock-on-lip operation at Mach number 2.0 with 25 degree half-angle spike. The cowling was attached to the fuselage through the boundary-layer plow and served as part of the fuselage boundary-layer diverter system. The performance of the half-cone inlet was compared with that of a scoop-type inlet and a normal-wedge inlet on a maximum-thrust-minus-drag basis. The increase in pressure recovery obtained with the half-cone inlet over that obtained with the reference inlets offset the slightly higher drags observed over the Mach number range for the half-cone so that the performance of this configuration was equal to that of the other inlets at Mach number 2.0 and was slightly superior at the lower Mach numbers. For a particular configuration, a peak pressure recovery of 0.879 was obtained at Mach number 2.0, zero angle of attack, and 4-percent throat bleed; the subcritical stability was 16 percent. Use of a fuselage-mounted boundary-layer splitter plate ahead of the inlet did not improve the stability. Subcritical distortion values were below 10 percent for all configurations. (author)

Measurement and Analysis of the Noise Radiated by Low Mach Number Centrifugal Blowers.

NASA Astrophysics Data System (ADS)

Yeager, David Marvin

An investigation was performed of the broad band, aerodynamically generated noise in low tip-speed Mach number, centrifugal air moving devices. An interdisciplinary experimental approach was taken which involved investigation of the aerodynamic and acoustic fields, and their mutual relationship. The noise generation process was studied using two experimental vehicles: (1) a scale model of a homologous family of centrifugal blowers typical of those used to cool computer and business equipment, and (2) a single blade from a centrifugal blower impeller placed in a known, controllable flow field. The radiation characteristics of the model blower were investigated by measuring the acoustic intensity distribution near the blower inlet and comparing it with the intensity near the inlet to an axial flow fan. Results showed that the centrifugal blower is a distributed, random noise source, unlike an axial fan which exhibited the effects of a coherent, interacting source distribution. Aerodynamic studies of the flow field in the inlet and at the discharge to the rotating impeller were used to assess the mean flow distribution through the impeller blade channels and to identify regions of excessive turbulence near the rotating blade row. Both circumferential and spanwise mean flow nonuniformities were identified along with a region of increased turbulence just downstream of the scroll cutoff. The fluid incidence angle, normally taken as an indicator of blower performance, was estimated from mean flow data as deviating considerably from an ideal impeller design. An investigation of the noise radiated from the single, isolated airfoil was performed using modern correlation and spectral analysis techniques. Radiation from the single blade in flow was characterized using newly developed expressions for the correlation area and the dipole source strength per unit area, and from the relationship between the blade surface pressure and the incident turbulent flow field. Results showed that radiation from the single blade was dominated by the effects of the incident turbulence. Normalized correlations areas of approximately 25% were measured at low frequencies. While the noise generation was more efficient at the trailing edge of the isolated blade, more noise was radiated from the region near the leading edge.

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.

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.

Background-Oriented Schlieren used in a hypersonic inlet test at NASA GRC

NASA Technical Reports Server (NTRS)

Clem, Michelle; Woike, Mark; Saunders, John

2016-01-01

Background Oriented Schlieren (BOS) is a derivative of the classical schlieren technology, which is used to visualize density gradients, such as shock wave structures in a wind tunnel. Changes in refractive index resulting from density gradients cause light rays to bend, resulting in apparent motion of a random background pattern. The apparent motion of the pattern is determined using cross-correlation algorithms (between no-flow and with-flow image pairs) producing a schlieren-like image. One advantage of BOS is its simplified setup which enables a larger field-of-view (FOV) than traditional schlieren systems. In the present study, BOS was implemented into the Combined Cycle Engine Large-Scale Inlet Mode Transition Experiment (CCE LIMX) in the 10x10 Supersonic Wind Tunnel at NASA Glenn Research Center. The model hardware for the CCE LIMX accommodates a fully integrated turbine based combined cycle propulsion system. To date, inlet mode transition between turbine and ramjet operation has been successfully demonstrated. High-speed BOS was used to visualize the behavior of the flow structures shock waves during unsteady inlet unstarts, a phenomenon known as buzz. Transient video images of inlet buzz were recorded for both the ramjet flow path (high speed inlet) and turbine flow path (low speed inlet). To understand the stability limits of the inlet, operation was pushed to the point of unstart and buzz. BOS was implemented in order to view both inlets simultaneously, since the required FOV was beyond the capability of the current traditional schlieren system. An example of BOS data (Images 1-6) capturing inlet buzz are presented.

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.

The Effect of Break Edge Configuration on the Aerodynamics of Anti-Ice Jet Flow

NASA Astrophysics Data System (ADS)

Tatar, V.; Yildizay, H.; Aras, H.

2015-05-01

One of the components of a turboprop gas turbine engine is the Front Bearing Structure (FBS) which leads air into the compressor. FBS directly encounters with ambient air, as a consequence ice accretion may occur on its static vanes. There are several aerodynamic parameters which should be considered in the design of anti-icing system of FBS, such as diameter, position, exit angle of discharge holes, etc. This research focuses on the effects of break edge configuration over anti-ice jet flow. Break edge operation is a process which is applied to the hole in order to avoid sharp edges which cause high stress concentration. Numerical analyses and flow visualization test have been conducted. Four different break edge configurations were used for this investigation; without break edge, 0.35xD, 74xD, 0.87xD. Three mainstream flow conditions at the inlet of the channel are defined; 10m/s, 20 m/s and 40 m/s. Shear stresses are extracted from numerical analyses near the trailing edge of pressure surface where ice may occur under icing conditions. A specific flow visualization method was used for the experimental study. Vane surface near the trailing edge was dyed and thinner was injected into anti-ice jet flow in order to remove dye from the vane surface. Hence, film effect on the surface could be computed for each testing condition. Thickness of the dye removal area of each case was examined. The results show noticeable effects of break edge operation on jet flow, and the air film effectiveness decreases when mainstream inlet velocity decreases.

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.

An experimental investigation of two large annular diffusers with swirling and distorted inflow

NASA Technical Reports Server (NTRS)

Eckert, W. T.; Johnston, J. P.; Simons, T. D.; Mort, K. W.; Page, V. R.

1980-01-01

Two annular diffusers downstream of a nacelle-mounted fan were tested for aerodynamic performance, measured in terms of two static pressure recovery parameters (one near the diffuser exit plane and one about three diameters downstream in the settling duct) in the presence of several inflow conditions. The two diffusers each had an inlet diameter of 1.84 m, an area ratio of 2.3, and an equivalent cone angle of 11.5, but were distinguished by centerbodies of different lengths. The dependence of diffuser performance on various combinations of swirling, radially distorted, and/or azimuthally distorted inflow was examined. Swirling flow and distortions in the axial velocity profile in the annulus upstream of the diffuser inlet were caused by the intrinsic flow patterns downstream of a fan in a duct and by artificial intensification of the distortions. Azimuthal distortions or defects were generated by the addition of four artificial devices (screens and fences). Pressure recovery data indicated beneficial effects of both radial distortion (for a limited range of distortion levels) and inflow swirl. Small amounts of azimuthal distortion created by the artificial devices produced only small effects on diffuser performance. A large artificial distortion device was required to produce enough azimuthal flow distortion to significantly degrade the diffuser static pressure recovery.

Sensitivity of shock boundary-layer interactions to weak geometric perturbations

NASA Astrophysics Data System (ADS)

Kim, Ji Hoon; Eaton, John K.

2016-11-01

Shock-boundary layer interactions can be sensitive to small changes in the inlet flow and boundary conditions. Robust computational models must capture this sensitivity, and validation of such models requires a suitable experimental database with well-defined inlet and boundary conditions. To that end, the purpose of this experiment is to systematically document the effects of small geometric perturbations on a SBLI flow to investigate the flow physics and establish an experimental dataset tailored for CFD validation. The facility used is a Mach 2.1, continuous operation wind tunnel. The SBLI is generated using a compression wedge; the region of interest is the resulting reflected shock SBLI. The geometric perturbations, which are small spanwise rectangular prisms, are introduced ahead of the compression ramp on the opposite wall. PIV is used to study the SBLI for 40 different perturbation geometries. Results show that the dominant effect of the perturbations is a global shift of the SBLI itself. In addition, the bumps introduce weaker shocks of varying strength and angles, depending on the bump height and location. Various scalar validation metrics, including a measure of shock unsteadiness, and their uncertainties are also computed to better facilitate CFD validation. Ji Hoon Kim is supported by an OTR Stanford Graduate Fellowship.

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.

Parametric study of the lubrication of thrust loaded 120-mm bore ball bearings to 3 million DN

NASA Technical Reports Server (NTRS)

Signer, H.; Bamberger, E. N.; Zaretsky, E. V.

1973-01-01

A parametric study was performed with 120-mm bore angular-contact ball bearings under varying thrust loads, bearing and lubricant temperatures, and cooling and lubricant flow rates. Contact angles were nominally 20 and 24 deg with bearing speeds to 3 million DN. Endurance tests were run at 3 million DN and a temperature of 492 K (425 F) with 10 bearings having a nominal 24 deg contact angle at a thrust load of 22241 N (5000 lb). Bearing operating temperature, differences in temperatures between the inner and outer races, and bearing power consumption can be tuned to any desirable operating requirement by varying 4 parameters. These parameters are outer-race cooling, inner-race cooling, lubricant flow to the inner race, and oil inlet temperature. Preliminary endurance tests at 3 million DN and 492 K (425 F) indicate that long term bearing operation can be achieved with a high degree of reliability.

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

Inlet Performance Characteristics from Wind-Tunnel Tests of a 0.10-Scale Air-Induction System Model of the YF-108A Airplane at Mach Numbers of 2.50, 2.76, and 3.00

NASA Technical Reports Server (NTRS)

Blackaby, James R.; Lyman, E. Gene; Altermann, John A., III

1959-01-01

Inlet-performance and external-drag-coefficient characteristics are presented without analysis. Effects are shown of variations of fuselage boundary-layer diverter profile, bleed-surface porosity, bleed-exit area, and inlet ramp, and lip angle.

An aerodynamic investigation of two 1.83-meter-diameter fan systems designed to drive a subsonic wind tunnel

NASA Technical Reports Server (NTRS)

Page, V. R.; Eckert, W. T.; Mort, K. W.

1977-01-01

An experimental, aerodynamic investigation was made of two 1.83 m diameter fan systems which are being considered for the repowered drive section of the 40- by 80-foot wind tunnel at NASA Ames Research Center. One system was low speed, the other was high speed. The low speed fan was tested at various stagger angles from 32.9 deg to 62.9 deg. At a fan blade stagger angle of 40.8 deg and operating at a tip speed of 1155 m/sec, the low speed fan developed 207.3 m of head. The high speed fan had a design blade stagger angle of 56.2 deg and was tested at this stagger angle only. The high speed fan operating at 191.5 m/sec developed 207.3 m of head. Radial distributions of static pressure coefficients, total pressure coefficients, and angles of swirl are presented. Radial surveys were conducted at four azimuth locations in front of the fan, and repeated downstream of the fan. Data were taken for various flow control devices and for two inlet contraction lengths.

Vane Separation Control in a Linear Cascade with Area Expansion using AC DBD Plasma Actuators

NASA Astrophysics Data System (ADS)

Kleven, Christopher; Corke, Thomas

2013-11-01

Experiments are presented on the use of AC dielectric barrier discharge (DBD) plasma actuators to prevent flow separation on vanes in a linear cascade with area expansion. The inlet Mach number to the cascade ranged from 0.3 to 0.5, and the vane chord Reynolds numbers ranged from 0 . 9 ×106 to 1 . 5 ×106 . Three cascade designs with different amounts of area expansion, providing different degrees of adverse pressure gradients, were examined. Surface flow visualization revealed a 3-D separation bubble with strong recirculation that formed on the suction side of the vanes. The pattern agreed well with CFD simulations. Plasma actuators were placed on the suction sides of the vanes, just upstream of the flow separation location. Quantitative measurements were performed in the wakes of the vanes using a 5-hole Pitot probe. The measurements were used to determine the effect of the plasma actuator separation control on the pressure loss coefficient, and flow turning angle through the cascades. Overall, the plasma actuators separation control increased the velocity magnitude and dynamic pressure in the passage between the vanes, resulted in a more spanwise-uniform flow turning angle in the vane passage, and significantly lowered the loss coefficient compared to the baseline.

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.

Experimental investigation of inlet-combustor isolators for a dual-mode scramjet at a Mach number of 4

NASA Technical Reports Server (NTRS)

Emami, Saied; Trexler, Carl A.; Auslender, Aaron H.; Weidner, John P.

1995-01-01

This report details experimentally derived operational characteristics of numerous two-dimensional planar inlet-combustor isolator configurations at a Mach number of 4. Variations in geometry included (1) inlet cowl length; (2) inlet cowl rotation angle; (3) isolator length; and (4) utilization of a rearward-facing isolator step. To obtain inlet-isolator maximum pressure-rise data relevant to ramjet-engine combustion operation, configurations were mechanically back pressured. Results demonstrated that the combined inlet-isolator maximum back-pressure capability increases as a function of isolator length and contraction ratio, and that the initiation of unstart is nearly independent of inlet cowl length, inlet cowl contraction ratio, and mass capture. Additionally, data are presented quantifying the initiation of inlet unstarts and the corresponding unstart pressure levels.

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.

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.

Nonimaging radiant energy device

DOEpatents

Winston, Roland; Ning, Xiaohui

1993-01-01

A nonimaging radiant energy device may include a hyperbolically shaped reflective element with a radiant energy inlet and a radiant energy outlet. A convex lens is provided at the radiant energy inlet and a concave lens is provided at the radiant energy outlet. Due to the provision of the lenses and the shape of the walls of the reflective element, the radiant energy incident at the radiant energy inlet within a predetermined angle of acceptance is emitted from the radiant energy outlet exclusively within an acute exit angle. In another embodiment, the radiant energy device may include two interconnected hyperbolically shaped reflective elements with a respective convex lens being provided at each aperture of the device.

Nonimaging radiant energy device

DOEpatents

Winston, Roland; Ning, Xiaohui

1996-01-01

A nonimaging radiant energy device may include a hyperbolically shaped reflective element with a radiant energy inlet and a radiant energy outlet. A convex lens is provided at the radiant energy inlet and a concave lens is provided at the radiant energy outlet. Due to the provision of the lenses and the shape of the walls of the reflective element, the radiant energy incident at the radiant energy inlet within a predetermined angle of acceptance is emitted from the radiant energy outlet exclusively within an acute exit angle. In another embodiment, the radiant energy device may include two interconnected hyperbolically shaped reflective elements with a respective convex lens being provided at each aperture of the device.

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.

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.

Computer code for preliminary sizing analysis of axial-flow turbines

NASA Technical Reports Server (NTRS)

Glassman, Arthur J.

1992-01-01

This mean diameter flow analysis uses a stage average velocity diagram as the basis for the computational efficiency. Input design requirements include power or pressure ratio, flow rate, temperature, pressure, and rotative speed. Turbine designs are generated for any specified number of stages and for any of three types of velocity diagrams (symmetrical, zero exit swirl, or impulse) or for any specified stage swirl split. Exit turning vanes can be included in the design. The program output includes inlet and exit annulus dimensions, exit temperature and pressure, total and static efficiencies, flow angles, and last stage absolute and relative Mach numbers. An analysis is presented along with a description of the computer program input and output with sample cases. The analysis and code presented herein are modifications of those described in NASA-TN-D-6702. These modifications improve modeling rigor and extend code applicability.

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.

Ram-recovery Characteristics of NACA Submerged Inlets at High Subsonic Speeds

NASA Technical Reports Server (NTRS)

Hall, Charles F; Frank, Joseph L

1948-01-01

Results are presented of an experimental investigation of the characteristics of NACA submerged inlets on a model of a fighter airplane for Mach numbers from 0.30 to 0.875. The effects on the ram-recovery ratio at the inlets of Mach number, angle of attack, boundary-layer thickness on the fuselage, inlet location, and boundary-layer deflectors are shown. The data indicate only a slight decrease in ram-recovery ratio for the inlets ahead of or just behind the wing leading edge as Mach number increased, but showed large decreases at high Mach numbers for the inlets aft of the point of maximum thickness of the wing.

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.

Predicting broadband noise from a stator vane of a gas turbine engine

NASA Technical Reports Server (NTRS)

Hanson, Donald B. (Inventor)

2002-01-01

A computer-implemented model of fan section of a gas turbine engine accounts for the turbulence in the gas flow emanating from the rotor assembly and impinging upon an inlet to the stator vane cascade. The model allows for user-input variations in the sweep and/or lean angles for the stator vanes. The model determines the resulting acoustic response of the fan section as a function of the turbulence and the lean and/or sweep angles of the vanes. The model may be embodied in software that is rapidly executed in a computer. This way, an optimum arrangement in terms of fan noise reduction is quickly determined for the stator vane lean and sweep physical positioning in the fan section of a gas turbine engine.

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

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.

Atmospheric Disturbance Environment Definition

NASA Technical Reports Server (NTRS)

Tank, William G.

1994-01-01

Traditionally, the application of atmospheric disturbance data to airplane design problems has been the domain of the structures engineer. The primary concern in this case is the design of structural components sufficient to handle transient loads induced by the most severe atmospheric "gusts" that might be encountered. The concern has resulted in a considerable body of high altitude gust acceleration data obtained with VGH recorders (airplane velocity, V, vertical acceleration, G, altitude, H) on high-flying airplanes like the U-2 (Ehernberger and Love, 1975). However, the propulsion system designer is less concerned with the accelerations of the airplane than he is with the airflow entering the system's inlet. When the airplane encounters atmospheric turbulence it responds with transient fluctuations in pitch, yaw, and roll angles. These transients, together with fluctuations in the free-stream temperature and pressure will disrupt the total pressure, temperature, Mach number and angularity of the inlet flow. For the mixed compression inlet, the result is a disturbed throat Mach number and/or shock position, and in extreme cases an inlet unstart can occur (cf. Section 2.1). Interest in the effects of inlet unstart on the vehicle dynamics of large, supersonic airplanes is not new. Results published by NASA in 1962 of wind tunnel studies of the problem were used in support of the United States Supersonic Transport program (SST) (White, at aI, 1963). Such studies continued into the late 1970's. However, in spite of such interest, there never was developed an atmospheric disturbance database for inlet unstart analysis to compare with that available for the structures load analysis. Missing were data for the free-stream temperature and pressure disturbances that also contribute to the unStart problem.

Genetic algorithm optimization of a film cooling array on a modern turbine inlet vane

NASA Astrophysics Data System (ADS)

Johnson, Jamie J.

In response to the need for more advanced gas turbine cooling design methods that factor in the 3-D flowfield and heat transfer characteristics, this study involves the computational optimization of a pressure side film cooling array on a modern turbine inlet vane. Latin hypersquare sampling, genetic algorithm reproduction, and Reynolds-Averaged Navier Stokes (RANS) computational fluid dynamics (CFD) as an evaluation step are used to assess a total of 1,800 film cooling designs over 13 generations. The process was efficient due to the Leo CFD code's ability to estimate cooling mass flux at surface grid cells using a transpiration boundary condition, eliminating the need for remeshing between designs. The optimization resulted in a unique cooling design relative to the baseline with new injection angles, compound angles, cooling row patterns, hole sizes, a redistribution of cooling holes away from the over-cooled midspan to hot areas near the shroud, and a lower maximum surface temperature. To experimentally confirm relative design trends between the optimized and baseline designs, flat plate infrared thermography assessments were carried out at design flow conditions. Use of flat plate experiments to model vane pressure side cooling was justified through a conjugate heat transfer CFD comparison of the 3-D vane and flat plate which showed similar cooling performance trends at multiple span locations. The optimized flat plate model exhibited lower minimum surface temperatures at multiple span locations compared to the baseline. Overall, this work shows promise of optimizing film cooling to reduce design cycle time and save cooling mass flow in a gas turbine.

An experimental investigation of endwall profiling in a turbine vane cascade

NASA Technical Reports Server (NTRS)

Kopper, F. C.; Milano, R.; Vanco, M.

1980-01-01

Measurements of surface static pressures, flow total pressure loss, and exit air angle were obtained for two linear cascades to establish the effects of endwall profiling. Testing was conducted at an isentropic exit Mach number of 0.85. One cascade was fabricated with planar endwalls while the other had one planar and one profiled endwall. Both cascades utilized the same high pressure turbine inlet guide vane section. It was found that in terms of full passage loss the profiled endwall cascade has the superior performance. The secondary loss results obtained are reasonably well predicted by correlations developed from incompressible flow testing of similar configurations. Inviscid flow and boundary layer calculations are compared with the test data, and overall, the agreement is found to be good. Use of the results for design purposes is briefly discussed.

Flow solution on a dual-block grid around an airplane

NASA Technical Reports Server (NTRS)

Eriksson, Lars-Erik

1987-01-01

The compressible flow around a complex fighter-aircraft configuration (fuselage, cranked delta wing, canard, and inlet) is simulated numerically using a novel grid scheme and a finite-volume Euler solver. The patched dual-block grid is generated by an algebraic procedure based on transfinite interpolation, and the explicit Runge-Kutta time-stepping Euler solver is implemented with a high degree of vectorization on a Cyber 205 processor. Results are presented in extensive graphs and diagrams and characterized in detail. The concentration of grid points near the wing apex in the present scheme is shown to facilitate capture of the vortex generated by the leading edge at high angles of attack and modeling of its interaction with the canard wake.

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.

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.

Standardized performance tests of collectors of solar thermal energy: Prototype moderately concentrating grooved collectors

NASA Technical Reports Server (NTRS)

1976-01-01

Prototypes of moderately concentrating grooved collectors were tested with a solar simulator for varying inlet temperature, flux level, and incident angle. Collector performance is correlated in terms of inlet temperature and flux level.

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.

Wind tunnel investigation of three axisymmetric cowls of different lengths at Mach numbers from 0.60 to 0.92

NASA Technical Reports Server (NTRS)

Re, Richard J.; Abeyounis, William K.

1993-01-01

Pressure distributions on three inlets having different cowl lengths were obtained in the Langley 16-Foot Transonic Tunnel. The cowl diameter ratio (highlight diameter to maximum diameter) was 0.85 and the cowl length ratios (cowl length to maximum diameter) were 0.337, 0.439, and 0.547. The cowls had identical nondimensionalized (with respect to cowl length) external geometry and identical internal geometry. The internal contraction ratio (highlight area to throat area) was 1.250. The inlets had longitudinal rows of static pressure orifices on the top and bottom (external) surfaces and on the contraction (internal) and diffuser surfaces. The afterbody was cylindrical in shape, and its diameter was equal to the maximum diameter of the cowl. Depending on the cowl configuration and free-stream Mach number, the mass-flow ratio varied between 0.27 and 0.87 during the tests. Angle of attack varied from 0 to 4.1 deg at selected Mach numbers and mass-flow ratios, and the Reynolds number varied with the Mach number from 3.2x10(exp 6) to 4.2x10(exp 6) per foot.

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.

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.

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.

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.

Fuel injection and mixing systems and methods of using the same

DOEpatents

Mao, Chien-Pei; Short, John

2010-08-03

A fuel injection and mixing system is provided. The system includes an injector body having a fuel inlet and a fuel outlet, and defines a fuel flow path between the inlet and outlet. The fuel flow path may include a generally helical flow passage having an inlet end portion disposed proximate the fuel inlet of the injector body. The flow path also may include an expansion chamber downstream from and in fluid communication with the helical flow passage, as well as a fuel delivery device in fluid communication with the expansion chamber for delivering fuel. Heating means is also provided in thermal communication with the injector body. The heating means may be adapted and configured for maintaining the injector body at a predetermined temperature to heat fuel traversing the flow path. A method of preheating and delivering fuel is also provided.

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.

Comparison of Computational and Experimental Results for a Transonic Variable-Speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

NASA Technical Reports Server (NTRS)

Booth, David; Flegel, Ashlie

2015-01-01

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

Comparison of Computational and Experimental Results for a Transonic Variable-speed Power-Turbine Blade Operating with Low Inlet Turbulence Levels

NASA Technical Reports Server (NTRS)

Booth, David T.; Flegel, Ashlie B.

2015-01-01

A computational assessment of the aerodynamic performance of the midspan section of a variable-speed power-turbine blade is described. The computation comprises a periodic single blade that represents the 2-D Midspan section VSPT blade that was tested in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. A commercial, off-the-shelf (COTS) software package, Pointwise and CFD++, was used for the grid generation and RANS and URANS computations. The CFD code, which offers flexibility in terms of turbulence and transition modeling options, was assessed in terms of blade loading, loss, and turning against test data from the transonic tunnel. Simulations were assessed at positive and negative incidence angles that represent the turbine cruise and take-off design conditions. The results indicate that the secondary flow induced at the positive incidence cruise condition results in a highly loaded case and transitional flow on the blade is observed. The negative incidence take-off condition is unloaded and the flow is very two-dimensional. The computational results demonstrate the predictive capability of the gridding technique and COTS software for a linear transonic turbine blade cascade with large incidence angle variation.

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 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.

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).

Vertical/Short Takeoff and Landing Model in the 10- by 10-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1979-05-21

A technician checks a 0.25-scale engine model of a Vought Corporation V-530 engine in the test section of the 10- by 10-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Vought created a low-drag tandem-fan Vertical/Short and Takeoff and Landing (V/STOL) engine in the mid-1970s, designated as the V-530. The first fan on the tandem-fan engine was supplied with air through a traditional subsonic inlet, seen on the lower front of the engine. The air was exhausted through the nacelle during normal flight and directed down during takeoffs. The rear fan was supplied by the oval-shaped top inlet during all phases of the flight. The second fan exhausted its air through a rear vectorable nozzle. NASA Lewis and Vought partnered in the late 1970s to collect an array of inlet and nozzle design information on the tandem fan engines for the Navy. Vought created this .25-scale model of the V-530 for extensive testing in Lewis' 10- by 10-foot tunnel. During an early series of tests, the front fan was covered, and a turbofan simulator was used to supply air to the rear fan. The researchers then analyzed the performance of only the front fan inlet. During the final series of tests, the flow from the front fan was used to supply airflow to the rear fan. The researchers studied the inlet's recovery, distortion, and angle-of-attack limits over various flight conditions.

Computational Investigation of Novel Tip Leakage Mitigation Methods for High Pressure Turbine Blades

NASA Technical Reports Server (NTRS)

Ibrahim, Mounir; Gupta, Abhinav; Shyam, Vikram

2014-01-01

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this paper a computational study was conducted on the Energy Efficient Engine (EEE) High Pressure Turbine (HPT) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was validated against aerodynamic data acquired in the NASA Transonic Turbine Blade Cascade facility. The scope of the ongoing study is to computationally investigate how the tip leakage and overall blade losses are affected by (1) injection from the tip near the pressure side, (2) injection from the tip surface at the camber line, and (3) injection from the tip surface into the tip separation bubble. The objective is to identify the locations on the tip surface at which to place appropriately configured blowing keeping in mind the film cooling application of tip blowing holes. The validation was conducted at Reynolds numbers of 85,000, 343,000, and 685,000 and at engine realistic flow conditions. The coolant injection simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and utilized the SST turbulence model in FLUENT. The key parameters examined are the number of jets, jet angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30 deg, -30 deg, and 90 deg to the local free stream on the tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean camber line is also presented (30 deg toward suction side). In addition, the effect of a single hole is also compared to a flat tip with no injection. The results provide insight into tip flow control methods and can be used to guide further investigation into tip flow control. As noted in past research it is concluded that reducing leakage flow is not necessarily synonymous with reducing losses due to leakage.

Computational Investigation of Novel Tip Leakage Mitigation Methods for High Pressure Turbine Blades

NASA Technical Reports Server (NTRS)

Ibrahim, Mounir; Gupta, Abhinav; Shyam, Vikram

2014-01-01

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this paper a computational study was conducted on the EEE (Energy Efficient Engine) HPT (High Pressure Turbine) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was validated against aerodynamic data acquired in the NASA Transonic Turbine Blade Cascade facility. The scope of the ongoing study is to computationally investigate how the tip leakage and overall blade losses are affected by 1. injection from the tip near the pressure side, 2. injection from the tip surface at the camber line, and 3. injection from the tip surface into the tip separation bubble. The objective is to identify the locations on the tip surface at which to place appropriately configured blowing keeping in mind the film cooling application of tip blowing holes. The validation was conducted at Reynolds numbers of 85,000, 343,000 and 685,000 and at engine realistic flow conditions. The coolant injection simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and utilized the SST turbulence model in FLUENT. The key parameters examined are the number of jets, jet angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30 deg., -30 deg. and 90 deg. to the local free stream on the tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean camber line is also presented (30 degrees toward suction side). In addition, the effect of a single hole is also compared to a flat tip with no injection. The results provide insight into tip flow control methods and can be used to guide further investigation into tip flow control. As noted in past research it is concluded that reducing leakage flow is not necessarily synonymous with reducing losses due to leakage.

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.

The Development of an 8-inch by 8-inch Slotted Tunnel for Mach Numbers up to 1.28

NASA Technical Reports Server (NTRS)

Little, B. H., Jr.; Cubbage, James J., Jr.

1961-01-01

An 8-inch by 8-inch transonic tunnel model with test section slotted on two opposite walls was constructed in which particular emphasis -was given to the development of slot geometry, slot-flow reentry section, and short-diffuser configurations for good test-region flow and minimum total-pressure losses. Center-line static pressures through the test section, wall static pressures through the other parts of the tunnel, and total-pressure distributions at the inlet and exit stations of the diffuser were measured- With a slot length equal to two tunnel heights and 1/14 open-area-ratio slotted walls) a test region one tunnel height in length was obtained in which the deviation from the mean Mach number was less than +/- 0.01 up to Mach number 1.15. With 1/7 open-area-ratio slotted walls, a test region 0.84 tunnel heights in length with deviation less than +/- O.01 was obtained up to Mach number 1.26. Increasing the tunnel diffuser angle from 6.4 to 10 deg. increased pressure loss through the tunnel at Mach number 1.20 from 15 percent to 20 percent of the total pressure. The use of other diffusers with equivalent angles of 10 deg. but contoured so that the initial diffusion angle was less than 10 deg. and the final angle was 200 reduced the losses to as low as 16 percent. A method for changing the test-section Mach number rapidly by controlling the flow through a bypass line from the tunnel settling chamber to the slot-flow plenum chamber of the test section was very effective. The test-section Mach number was reduced approximately 5 percent in 1/8 second by bleeding into the test section a flow of air equal to 2 percent of the mainstream flow and 30 percent in 1/4 second with bleed flow equal to 10 percent of the mainstream flow. The rate of reduction was largely determined by the opening rate of the bleed-flow-control valve.

Simulation of personalised haemodynamics by various mounting positions of a prosthetic valve using computational fluid dynamics.

PubMed

Bongert, Markus; Geller, Marius; Pennekamp, Werner; Nicolas, Volkmar

2018-03-03

Diseases of the cardiovascular system account for nearly 42% of all deaths in the European Union. In Germany, approximately 12,000 patients receive surgical replacement of the aortic valve due to heart valve disease alone each year. A three-dimensional (3D) numerical model based on patient-specific anatomy derived from four-dimensional (4D) magnetic resonance imaging (MRI) data was developed to investigate preoperatively the flow-induced impact of mounting positions of aortic prosthetic valves to select the best orientation for individual patients. Systematic steady-state analysis of blood flow for different rotational mounting positions of the valve is only possible using a virtual patient model. A maximum velocity of 1 m/s was used as an inlet boundary condition, because the opening angle of the valve is at its largest at this velocity. For a comparative serial examination, it is important to define the standardised general requirements to avoid impacts other than the rotated implantation of the prosthetic aortic valve. In this study, a uniform velocity profile at the inlet for the inflow of the aortic valve and the real aortic anatomy were chosen for all simulations. An iterative process, with the weighted parameters flow resistance (1), shear stress (2) and velocity (3), was necessary to determine the best rotated orientation. Blood flow was optimal at a 45° rotation from the standard implantation orientation, which will offer a supply to the coronary arteries.

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.

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

Program to develop a performance and heat load prediction system for multistage turbines

NASA Technical Reports Server (NTRS)

Sharma, OM

1994-01-01

Flows in low-aspect ratio turbines, such as the SSME fuel turbine, are three dimensional and highly unsteady due to the relative motion of adjacent airfoil rows and the circumferential and spanwise gradients in total pressure and temperature, The systems used to design these machines, however, are based on the assumption that the flow is steady. The codes utilized in these design systems are calibrated against turbine rig and engine data through the use of empirical correlations and experience factors. For high aspect ratio turbines, these codes yield reasonably accurate estimates of flow and temperature distributions. However, future design trends will see lower aspect ratio (reduced number of parts) and higher inlet temperature which will result in increased three dimensionality and flow unsteadiness in turbines. Analysis of recently acquired data indicate that temperature streaks and secondary flows generated in combustors and up-stream airfoils can have a large impact on the time-averaged temperature and angle distributions in downstream airfoil rows.

Wall boundary layer development near the tip region of an IGV of an axial flow compressor

NASA Technical Reports Server (NTRS)

Lakshminarayana, B.; Sitaram, N.

1983-01-01

The annulus wall boundary layer inside the blade passage of the inlet guide vane (IGV) passage of a low-speed axial compressor stage was measured with a miniature five-hole probe. The three-dimensional velocity and pressure fields were measured at various axial and tangential locations. Limiting streamline angles and static pressures were also measured on the casing of the IGV passage. Strong secondary vorticity was developed. The data were analyzed and correlated with the existing velocity profile correlations. The end wall losses were also derived from these data.

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.

Fuel injector

DOEpatents

Lambeth, Malcolm David Dick [Bromley, GB

2001-02-27

A fuel injector comprises first and second housing parts, the first housing part being located within a bore or recess formed in the second housing part, the housing parts defining therebetween an inlet chamber, a delivery chamber axially spaced from the inlet chamber, and a filtration flow path interconnecting the inlet and delivery chambers to remove particulate contaminants from the flow of fuel therebetween.

Enhanced Performance of Streamline-Traced External-Compression Supersonic Inlets

NASA Technical Reports Server (NTRS)

Slater, John W.

2015-01-01

A computational design study was conducted to enhance the aerodynamic performance of streamline-traced, external-compression inlets for Mach 1.6. The current study explored a new parent flowfield for the streamline tracing and several variations of inlet design factors, including the axial displacement and angle of the subsonic cowl lip, the vertical placement of the engine axis, and the use of porous bleed in the subsonic diffuser. The performance was enhanced over that of an earlier streamline-traced inlet such as to increase the total pressure recovery and reduce total pressure distortion

A novel concept for subsonic inlet boundary-layer control

NASA Technical Reports Server (NTRS)

Miller, B. A.

1977-01-01

A self-bleeding method for boundary layer control is described and tested for a subsonic inlet designed to operate in the flowfield generated by high angles of attack. Naturally occurring surface static pressure gradients are used to remove the boundary layer from a separation-prone region of the inlet and to reinject it at a less critical location with a net performance gain. The results suggest that this self-bleeding method for boundary-layer control might be successfully applied to other inlets operating at extreme aerodynamic conditions.

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.

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.

Comparable ultrasound measurements of ten anatomical specimens of infant hip joints by the methods of Graf and Terjesen.

PubMed

Falliner, A; Hahne, H J; Hedderich, J; Brossmann, J; Hassenpflug, J

2004-04-01

To define which sonographic section planes relative to the acetabular inlet plane will produce analyzable images with the methods of Graf and Terjesen. Anatomical specimens of infant hip joints were investigated in a water bath using the methods of Graf and Terjesen. Acetabular position was varied in defined increments with respect to the ultrasound beam. The alpha angles and the femoral head coverage (FHC) were measured. To obtain images analyzable by the two methods, the ultrasound beam had to intersect with the acetabular inlet plane at defined angles. The acetabular notch had to be anteriorly rotated from the ultrasound beam plane by at least 20 degrees. Beam entry within a 50 degrees sector posterior to the perpendicular on the inlet plane resulted in analyzable images. The stepwise multiple linear regression analysis showed that alpha angles and FHC were much affected by the coronal-plane transducer tilt. The fact that caudal tilts of the transducer are associated with reduced alpha angles and FHC values should be kept in mind in clinical ultrasound investigations. It is recommended that the transducer should be put on the greater trochanter perpendicular to the transverse axis of the body.

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.

On the Physics of Flow Separation Along a Low Pressure Turbine Blade Under Unsteady Flow Conditions

NASA Technical Reports Server (NTRS)

Schobeiri, Meinhard T.; Ozturk, Burak; Ashpis, David E.

2005-01-01

The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flow conditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the physics of the separation phenomenon under periodic unsteady wake flow. Several physical mechanisms are discussed.

An aerodynamic design and numerical investigation of transonic centrifugal compressor stage

NASA Astrophysics Data System (ADS)

Yi, Weilin; Ji, Lucheng; Tian, Yong; Shao, Weiwei; Li, Weiwei; Xiao, Yunhan

2011-09-01

In the present paper, the design of a transonic centrifugal compressor stage with the inlet relative Mach number about 1.3 and detailed flow field investigation by three-dimensional CFD are described. Firstly the CFD program was validated by an experimental case. Then the preliminary aerodynamic design of stage completed through in-house one-dimensional code. Three types of impellers and two sets of stages were computed and analyzed. It can be found that the swept shape of leading edge has prominent influence on the performance and can enlarge the flow range. Similarly, the performance of the stage with swept impeller is better than others. The total pressure ratio and adiabatic efficiency of final geometry achieve 7:1 and 80% respectively. The vane diffuser with same airfoils along span increases attack angle at higher span, and the local flow structure and performance is deteriorated.

Experimental evaluation of a cooled radial-inflow turbine

NASA Technical Reports Server (NTRS)

Tirres, Lizet; Dicicco, L. D.; Nowlin, Brent C.

1993-01-01

Two 14.4 inch tip diameter rotors were installed and tested in the Small Engines Component Turbine Facility (SECTF) at NASA Lewis Research Center. The rotors, a solid and a cooled version of a radial-inflow turbine, were tested with a 15 vane stat or over a set of rotational speeds ranging from 80 to 120 percent design speed (17,500 to 21,500 rpm). The total-to-total stage pressure ratios ranged from 2.5 to 5.5. The data obtained at the equivalent conditions using the solid version of the rotor are presented with the cooled rotor data. A Reynolds number of 381,000 was maintained for both rotors, whose stages had a design mass flow of 4.0 lbm/sec, a design work level of 59.61 Btu/lbm, and a design efficiency of 87 percent. The results include mass flow data, turbine torque, turbine exit flow angles, stage efficiency, and rotor inlet and exit surveys.

Experimental Evaluation of a Cooled Radial-inflow Turbine

NASA Technical Reports Server (NTRS)

Tirres, Lizet; Dicicco, L. Danielle; Nowlin, Brent C.

1993-01-01

Two 14.4 inch tip diameter rotors were installed and tested in the Small Engines Component Turbine Facility (SECTF) at NASA Lewis Research Center. The rotors, a solid and a cooled version of a radial-inflow turbine, were tested with a 15 vane stat or over a set of rotational speeds ranging from 80 to 120 percent design speed (17,500 to 21,500 rpm). The total-to-total stage pressure ratios ranged from 2.5 to 5.5. The data obtained at the equivalent conditions using the solid version of the rotor are presented with the cooled rotor data. A Reynolds number of 381,000 was maintained for both rotors, whose stages had a design mass flow of 4.0 Ibm/sec, a design work level of 59.61 Btu/lbm, and a design efficiency of 87 percent. The results include mass flow data, turbine torque, turbine exit flow angles, stage efficiency, and rotor inlet and exit surveys.

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.

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.

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.

The design and development of transonic multistage compressors

NASA Technical Reports Server (NTRS)

Ball, C. L.; Steinke, R. J.; Newman, F. A.

1988-01-01

The development of the transonic multistage compressor is reviewed. Changing trends in design and performance parameters are noted. These changes are related to advances in compressor aerodynamics, computational fluid mechanics and other enabling technologies. The parameters normally given to the designer and those that need to be established during the design process are identified. Criteria and procedures used in the selection of these parameters are presented. The selection of tip speed, aerodynamic loading, flowpath geometry, incidence and deviation angles, blade/vane geometry, blade/vane solidity, stage reaction, aerodynamic blockage, inlet flow per unit annulus area, stage/overall velocity ratio, and aerodynamic losses are considered. Trends in these parameters both spanwise and axially through the machine are highlighted. The effects of flow mixing and methods for accounting for the mixing in the design process are discussed.

An Inlet Distortion Assessment During Aircraft Departures at High Angle of Attack for an F/A-18A Aircraft

NASA Technical Reports Server (NTRS)

Steenken, William G.; Williams, John G.; Yuhas, Andrew J.; Walsh, Kevin R.

1997-01-01

The F404-GE-400-powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the quality of inlet airflow during departed flight maneuvers, that is, during flight outside the normal maneuvering envelope where control surfaces have little or no effectiveness. Six nose-left and six nose-right departures were initiated at Mach numbers between 0.3 and 0.4 at an altitude of 35 kft. The entry yaw rates were approximately 40 to 90 deg/sec. Engine surges were encountered during three of the nose-left and one of the nose-right departures. Time-variant inlet-total-pressure distortion levels at the engine face did not significantly exceed those at maximum angle-of-attack and sideslip maneuvers during controlled flight. Surges caused by inlet distortion levels resulted from a combination of high levels of inlet distortion and rapid changes in aircraft position. These rapid changes indicate a combination of engine support and gyroscopic loads being applied to the engine structure that impact the aerodynamic stability of the compressor through changes in the rotor-to-case clearances. This document presents the slides from an oral presentation.

On the Physics of Flow Separation Along a Low Pressure Turbine Blade Under Unsteady Flow Conditions

NASA Technical Reports Server (NTRS)

Schobeiri, Meinhard T.; Ozturk, Burak; Ashpis, David E.

2003-01-01

The present study, which is the first of a series of investigations dealing with specific issues of low pressure turbine (LPT) boundary layer aerodynamics, is aimed at providing detailed unsteady boundary flow information to understand the underlying physics of the inception, onset, and extent of the separation zone. A detailed experimental study on the behavior of the separation zone on the suction surface of a highly loaded LPT-blade under periodic unsteady wake flow is presented. Experimental investigations were performed at Texas A&M Turbomachinery Performance and Flow Research Laboratory using a large-scale unsteady turbine cascade research facility with an integrated wake generator and test section unit. To account for a high flow deflection of LPT-cascades at design and off-design operating points, the entire wake generator and test section unit including the traversing system is designed to allow a precise angle adjustment of the cascade relative to the incoming flow. This is done by a hydraulic platform, which simultaneously lifts and rotates the wake generator and test section unit. The unit is then attached to the tunnel exit nozzle with an angular accuracy of better than 0.05 , which is measured electronically. Utilizing a Reynolds number of 110,000 based on the blade suction surface length and the exit velocity, one steady and two different unsteady inlet flowconditions with the corresponding passing frequencies, wake velocities and turbulence intensities are investigated using hot-wire anemometry. In addition to the unsteady boundary layer measurements, blade surface pressure measurements were performed at Re=50,000, 75,000, 100,000, and 125,000 at one steady and two periodic unsteady inlet flow conditions. Detailed unsteady boundary layer measurement identifies the onset and extent of the separation zone as well as its behavior under unsteady wake flow. The results presented in ensemble-averaged and contour plot forms contribute to understanding the physics of the separation phenomenon under periodic unsteady wake flow. Several physical mechanisms are discussed.

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.

Pore-scale modeling of wettability effects on CO2-brine displacement during geological storage

NASA Astrophysics Data System (ADS)

Basirat, Farzad; Yang, Zhibing; Niemi, Auli

2017-11-01

Wetting properties of reservoir rocks and caprocks can vary significantly, and they strongly influence geological storage of carbon dioxide in deep saline aquifers, during which CO2 is supposed to displace the resident brine and to become permanently trapped. Fundamental understanding of the effect of wettability on CO2-brine displacement is thus important for improving storage efficiency and security. In this study, we investigate the influence of wetting properties on two-phase flow of CO2 and brine at the pore scale. A numerical model based on the phase field method is implemented to simulate the two-phase flow of CO2-brine in a realistic pore geometry. Our focus is to study the pore-scale fluid-fluid displacement mechanisms under different wetting conditions and to quantify the effect of wettability on macroscopic parameters such as residual brine saturation, capillary pressure, relative permeability, and specific interfacial area. Our simulation results confirm that both the trapped wetting phase saturation and the normalized interfacial area increase with decreasing contact angle. However, the wetting condition does not appear to influence the CO2 breakthrough time and saturation. We also show that the macroscopic capillary pressures based on the pressure difference between inlet and outlet can differ significantly from the phase averaging capillary pressures for all contact angles when the capillary number is high (log Ca > -5). This indicates that the inlet-outlet pressure difference may not be a good measure of the continuum-scale capillary pressure. In addition, the results show that the relative permeability of CO2 can be significantly lower in strongly water-wet conditions than in the intermediate-wet conditions.

Partitioning dynamics of unsaturated flows in fractured porous media: Laboratory studies and three-dimensional multi-scale smoothed particle hydrodynamics simulations of gravity-driven flow in fractures

NASA Astrophysics Data System (ADS)

Kordilla, J.; Bresinsky, L. T.; Shigorina, E.; Noffz, T.; Dentz, M.; Sauter, M.; Tartakovsky, A. M.

2017-12-01

Preferential flow dynamics in unsaturated fractures remain a challenging topic on various scales. On pore- and fracture-scales the highly erratic gravity-driven flow dynamics often provoke a strong deviation from classical volume-effective approaches. Against the common notion that flow in fractures (or macropores) can only occur under equilibrium conditions, i.e., if the surrounding porous matrix is fully saturated and capillary pressures are high enough to allow filling of the fracture void space, arrival times suggest the existence of rapid preferential flow along fractures, fracture networks, and fault zones, even if the matrix is not fully saturated. Modeling such flows requires efficient numerical techniques to cover various flow-relevant physics, such as surface tension, static and dynamic contact angles, free-surface (multi-phase) interface dynamics, and formation of singularities. Here we demonstrate the importance of such flow modes on the partitioning dynamics at simple fracture intersections, with a combination of laboratory experiments, analytical solutions and numerical simulations using our newly developed massively parallel smoothed particle hydrodynamics (SPH) code. Flow modes heavily influence the "bypass" behavior of water flowing along a fracture junction. Flows favoring the formation of droplets exhibit a much stronger bypass capacity compared to rivulet flows, where nearly the whole fluid mass is initially stored within the horizontal fracture. This behavior is demonstrated for a multi-inlet laboratory setup where the inlet-specific flow rate is chosen so that either a droplet or rivulet flow persists. The effect of fluid buffering within the horizontal fracture is presented in terms of dimensionless fracture inflow so that characteristic scaling regimes can be recovered. For both cases (rivulets and droplets), flow within the horizontal fracture transitions into a Washburn regime until a critical threshold is reached and the bypass efficiency increases. For rivulet flows, the initial filling of the horizontal fracture is described by classical plug flow. Meanwhile, for droplet flows, a size-dependent partitioning behavior is observed, and the filling of the fracture takes longer.

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.

Thrust modulation methods for a subsonic V/STOL aircraft

NASA Technical Reports Server (NTRS)

Woollett, R. R.

1981-01-01

Low speed wind tunnel tests were conducted to assess four methods for attaining thrust modulation for V/STOL aircraft. The four methods were: (1) fan speed change, (2) fan nozzle exit area change, (3) variable pitch rotor (VPR) fan, and (4) variable inlet guide vanes (VIGV). The interrelationships between inlet and thrust modulation system were also investigated using a double slotted inlet and thick lip inlet. Results can be summarized as: (1) the VPR and VIGV systems were the most promising, (2) changes in blade angle to obtain changes in fan thrust have significant implications for the inlet, and (3) both systems attained required level of thrust with acceptable levels of fan blade stress.

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.

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.

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.

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.

Enhanced Performance of Streamline-Traced External-Compression Supersonic Inlets

NASA Technical Reports Server (NTRS)

Slater, John W.

2015-01-01

A computational design study was conducted to enhance the aerodynamic performance of streamline-traced, external-compression inlets for Mach 1.6. Compared to traditional external-compression, two-dimensional and axisymmetric inlets, streamline-traced inlets promise reduced cowl wave drag and sonic boom, but at the expense of reduced total pressure recovery and increased total pressure distortion. The current study explored a new parent flowfield for the streamline tracing and several variations of inlet design factors, including the axial displacement and angle of the subsonic cowl lip, the vertical placement of the engine axis, and the use of porous bleed in the subsonic diffuser. The performance was enhanced over that of an earlier streamline-traced inlet such as to increase the total pressure recovery and reduce total pressure distortion.

Pressure activated stability-bypass-control valves to increase the stable airflow range of a Mach 2.5 inlet with 40 percent internal contraction

NASA Technical Reports Server (NTRS)

Mitchell, G. A.; Sanders, B. W.

1974-01-01

The throat of a Mach 2.5 inlet with a coldpipe termination was fitted with a stability-bypass system. The inlet stable airflow range provided by various stability-bypass entrance configurations in alternate combination with several stability-bypass exit controls was determined for both steady-state conditions and internal transient pulses. Transient results were also obtained for the inlet with a choke point at the diffuser exit. Instart angles of attack were determined for the various stability-bypass entrance configurations. The response of the inlet-coldpipe system to internal and external oscillating disturbances was determined. Poppet valves at the stability-bypass exit provided an inlet stable airflow range of 28 percent or greater at all static and transient conditions.

Comparison of multifrequency acoustic and in situ measurements of zooplankton abundances in Knight Inlet, British Columbia.

PubMed

Trevorrow, Mark V; Mackas, David L; Benfield, Mark C

2005-06-01

An investigation of midwater zooplankton aggregations in a coastal fjord was conducted in November 2002. This study focused on quantitative comparisons between a calibrated, three-frequency (38, 120, and 200 kHz) vessel-based echo-sounder, a multinet towed zooplankton sampler (BIONESS), and a high-resolution underwater camera (ZOOVIS). Daytime layers of euphausiids and amphipods near 70-90-m depth were observed in lower parts of the inlet, especially concentrated by tidal flows around a sill. Quantitative backscatter measurements of euphausiids and amphipods, combined with in situ size and abundance estimates, and using an assumed tilt-angle distribution, were in agreement with averaged fluid-cylinder scattering models produced by Stanton and Chu [ICES J. Mar. Sci. 57, 793-807, (2000)]. Acoustic measurements of physonect siphonophores in the upper inlet were found to have a strong 38-kHz scattering strength, in agreement with a damped bubble scattering model using a diameter of 0.4 mm. In relatively dense euphausiid layers, ZOOVIS abundance estimates were found to be a factor of 2 to 4 higher than the acoustic estimates, potentially due to deviations from assumed euphausiid orientation. Nocturnal near-surface euphausiid scattering exhibited a strong (15 dB) and rapid (seconds) sensitivity to vessel lights, interpreted as due to changing animal orientation.

Comparison of multifrequency acoustic and in situ measurements of zooplankton abundances in Knight Inlet, British Columbia

NASA Astrophysics Data System (ADS)

Trevorrow, Mark V.; Mackas, David L.; Benfield, Mark C.

2005-06-01

An investigation of midwater zooplankton aggregations in a coastal fjord was conducted in November 2002. This study focused on quantitative comparisons between a calibrated, three-frequency (38, 120, and 200 kHz) vessel-based echo-sounder, a multinet towed zooplankton sampler (BIONESS), and a high-resolution underwater camera (ZOOVIS). Daytime layers of euphausiids and amphipods near 70-90-m depth were observed in lower parts of the inlet, especially concentrated by tidal flows around a sill. Quantitative backscatter measurements of euphausiids and amphipods, combined with in situ size and abundance estimates, and using an assumed tilt-angle distribution, were in agreement with averaged fluid-cylinder scattering models produced by Stanton and Chu [ICES J. Mar. Sci. 57, 793-807, (2000)]. Acoustic measurements of physonect siphonophores in the upper inlet were found to have a strong 38-kHz scattering strength, in agreement with a damped bubble scattering model using a diameter of 0.4 mm. In relatively dense euphausiid layers, ZOOVIS abundance estimates were found to be a factor of 2 to 4 higher than the acoustic estimates, potentially due to deviations from assumed euphausiid orientation. Nocturnal near-surface euphausiid scattering exhibited a strong (15 dB) and rapid (seconds) sensitivity to vessel lights, interpreted as due to changing animal orientation. .

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.

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.

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.

Turbulent flow separation in three-dimensional asymmetric diffusers

NASA Astrophysics Data System (ADS)

Jeyapaul, Elbert

2011-12-01

Turbulent three-dimensional flow separation is more complicated than 2-D. The physics of the flow is not well understood. Turbulent flow separation is nearly independent of the Reynolds number, and separation in 3-D occurs at singular points and along convergence lines emanating from these points. Most of the engineering turbulence research is driven by the need to gain knowledge of the flow field that can be used to improve modeling predictions. This work is motivated by the need for a detailed study of 3-D separation in asymmetric diffusers, to understand the separation phenomena using eddy-resolving simulation methods, assess the predictability of existing RANS turbulence models and propose modeling improvements. The Cherry diffuser has been used as a benchmark. All existing linear eddy-viscosity RANS models k--o SST,k--epsilon and v2- f fail in predicting such flows, predicting separation on the wrong side. The geometry has a doubly-sloped wall, with the other two walls orthogonal to each other and aligned with the diffuser inlet giving the diffuser an asymmetry. The top and side flare angles are different and this gives rise to different pressure gradient in each transverse direction. Eddyresolving simulations using the Scale adaptive simulation (SAS) and Large Eddy Simulation (LES) method have been used to predict separation in benchmark diffuser and validated. A series of diffusers with the same configuration have been generated, each having the same streamwise pressure gradient and parametrized only by the inlet aspect ratio. The RANS models were put to test and the flow physics explored using SAS-generated flow field. The RANS model indicate a transition in separation surface from top sloped wall to the side sloped wall at an inlet aspect ratio much lower than observed in LES results. This over-sensitivity of RANS models to transverse pressure gradients is due to lack of anisotropy in the linear Reynolds stress formulation. The complexity of the flow separation is due to effects of lateral straining, streamline curvature, secondary flow of second kind, transverse pressure gradient on turbulence. Resolving these effects is possible with anisotropy turbulence models as the Explicit Algebraic Reynolds stress model (EARSM). This model has provided accurate prediction of streamwise and transverse velocity, however the wall pressure is under predicted. An improved EARSM model is developed by correcting the coefficients, which predicts a more accurate wall pressure. There exists scope for improvement of this model, by including convective effects and dynamics of velocity gradient invariants.

Single-stage experimental evaluation of tandem-airfoil rotor and stator blading for compressors, part 8

NASA Technical Reports Server (NTRS)

Brent, J. A.; Clemmons, D. R.

1974-01-01

An experimental investigation was conducted with an 0.8 hub/tip ratio, single-stage, axial flow compressor to determine the potential of tandem-airfoil blading for improving the efficiency and stable operating range of compressor stages. The investigation included testing of a baseline stage with single-airfoil blading and two tandem-blade stages. The overall performance of the baseline stage and the tandem-blade stage with a 20-80% loading split was considerably below the design prediction. The other tandem-blade stage, which had a rotor with a 50-50% loading split, came within 4.5% of the design pressure rise (delta P(bar)/P(bar) sub 1) and matched the design stage efficiency. The baseline stage with single-airfoil blading, which was designed to account for the actual rotor inlet velocity profile and the effects of axial velocity ratio and secondary flow, achieved the design predicted performance. The corresponding tandem-blade stage (50-50% loading split in both blade rows) slightly exceeded the design pressure rise but was 1.5 percentage points low in efficiency. The tandem rotors tested during both phases demonstrated higher pressure rise and efficiency than the corresponding single-airfoil rotor, with identical inlet and exit airfoil angles.

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.