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.

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.

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

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

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.

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.

Viscous analyses for flow through subsonic and supersonic intakes

NASA Technical Reports Server (NTRS)

Povinelli, Louis A.; Towne, Charles E.

1986-01-01

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

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

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2014-01-01

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

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

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2014-01-01

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

Effects of installation caused flow distortion on noise from a fan designed for turbofan engines

NASA Technical Reports Server (NTRS)

Povinelli, F. P.; Dittmar, J. H.; Woodward, R. P.

1972-01-01

Far-field noise measurements were taken for three different installations of essentially the same fan. The installation with the most uniform inlet flow resulted in fan-blade-passage tone sound pressure levels more than 10 dB lower than the installation with more nonuniform inflow. Perceived noise levels were computed for the various installations and compared. Some measurements of inlet flow distortion were made and used in a blade-passage noise generation theory to predict the effects of distortion on noise. Good agreement was obtained between the prediction and the measured effect. Possible origins of the distortion were identified by observation of tuft action in the vicinity of the inlet.

Internal flow characteristics of a multistage compressor with inlet pressure distortion. [J85-13 turbojet engine studies

NASA Technical Reports Server (NTRS)

Debogdan, C. E.; Moss, J. E., Jr.; Braithwaite, W. M.

1977-01-01

The measured distribution of compressor interstage pressures and temperatures resulting from a 180 deg inlet-total-pressure distortion for a J85-13 turbojet engine is reported. Extensive inner stage instrumentation combined with stepwise rotation of the inlet distortion gave data of high circumferential resolution. The steady-state pressures and temperatures along with the amplitude, extent, and location of the distorted areas are given. Data for 80, 90, and 100 percent of rotor design speed are compared with clean (undistorted) inlet flow conditions to show pressure and temperature behavior within the compressor. Both overall and stagewise compressor performances vary only slightly when clean and distorted inlet conditions are compared. Total and static pressure distortions increase in amplitude in the first few stages of the compressor and then attenuate fairly uniformly to zero at the discharge. Total-temperature distortion induced by the pressure distortion reached a maximum amplitude by the first two stages and decayed only a little through the rest of the compressor. Distortion amplitude tended to peak in line with the screen edges, and, except for total and static pressure in the tip zone, there was little swirl in the axial direction.

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

NASA Technical Reports Server (NTRS)

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

1986-01-01

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

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.

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.

Effects of Distortion on Mass Flow Plug Calibration

NASA Technical Reports Server (NTRS)

Sasson, Jonathan; Davis, David O.; Barnhart, Paul J.

2015-01-01

A numerical, and experimental investigation to study the effects of flow distortion on a Mass Flow Plug (MFP) used to control and measure mass-flow during an inlet test has been conducted. The MFP was first calibrated using the WIND-US flow solver for uniform (undistorted) inflow conditions. These results are shown to compare favorably with an experimental calibration under similar conditions. The effects of distortion were investigated by imposing distorted flow conditions taken from an actual inlet test to the inflow plane of the numerical simulation. The computational fluid dynamic (CFD) based distortion study only showed the general trend in mass flow rate. The study used only total pressure as the upstream boundary condition, which was not enough to define the flow. A better simulation requires knowledge of the turbulence structure and a specific distortion pattern over a range of plug positions. It is recommended that future distortion studies utilize a rake with at least the same amount of pitot tubes as the AIP rake.

Estimating maximum instantaneous distortion from inlet total pressure rms and PSD measurements. [Root Mean Square and Power Spectral Density methods

NASA Technical Reports Server (NTRS)

Melick, H. C., Jr.; Ybarra, A. H.; Bencze, D. P.

1975-01-01

An inexpensive method is developed to determine the extreme values of instantaneous inlet distortion. This method also provides insight into the basic mechanics of unsteady inlet flow and the associated engine reaction. The analysis is based on fundamental fluid dynamics and statistical methods to provide an understanding of the turbulent inlet flow and quantitatively relate the rms level and power spectral density (PSD) function of the measured time variant total pressure fluctuations to the strength and size of the low pressure regions. The most probable extreme value of the instantaneous distortion is then synthesized from this information in conjunction with the steady state distortion. Results of the analysis show the extreme values to be dependent upon the steady state distortion, the measured turbulence rms level and PSD function, the time on point, and the engine response characteristics. Analytical projections of instantaneous distortion are presented and compared with data obtained by a conventional, highly time correlated, 40 probe instantaneous pressure measurement system.

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.

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.

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

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2014-01-01

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

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.

Effects of Inlet Distortion on Aeromechanical Stability of a Forward-Swept High-Speed Fan

NASA Technical Reports Server (NTRS)

Herrick, Gregory P.

2011-01-01

Concerns regarding noise, propulsive efficiency, and fuel burn are inspiring aircraft designs wherein the propulsive turbomachines are partially (or fully) embedded within the airframe; such designs present serious concerns with regard to aerodynamic and aeromechanic performance of the compression system in response to inlet distortion. Separately, a forward-swept high-speed fan was developed to address noise concerns of modern podded turbofans; however this fan encounters aeroelastic instability (flutter) as it approaches stall. A three-dimensional, unsteady, Navier-Stokes computational fluid dynamics code is applied to analyze and corroborate fan performance with clean inlet flow. This code, already validated in its application to assess aerodynamic damping of vibrating blades at various flow conditions, is modified and then applied in a computational study to preliminarily assess the effects of inlet distortion on aeroelastic stability of the fan. Computational engineering application and implementation issues are discussed, followed by an investigation into the aeroelastic behavior of the fan with clean and distorted inlets.

Some observations of the effects of radial distortions on performance of a transonic rotating blade row

NASA Technical Reports Server (NTRS)

Sandercock, D. M.; Sanger, N. L.

1974-01-01

A single rotating blade row was tested with two magnitudes of tip radial distortion and two magnitudes of hub radial distortion imposed on the inlet flow. The rotor was about 50 centimeters (20 in.) in diameter and had a design operating tip speed of approximately 420 meters per second (1380 ft/sec). Overall performance at 60, 80, and 100 percent of equivalent design speed generally showed a decrease (compared to undistorted flow) in rotor stall margin with tip radial distortion but no change, or a slight increase, in rotor stall margin with hub radial distortion. At design speed there was a decrease in rotor overall total pressure ratio and choke flow with all inlet flow distortions. Radial distributions of blade element parameters are presented for selected operating conditions at design speed.

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

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

Blade row dynamic digital compressor program. Volume 1: J85 clean inlet flow and parallel compressor models

NASA Technical Reports Server (NTRS)

Tesch, W. A.; Steenken, W. G.

1976-01-01

The results are presented of a one-dimensional dynamic digital blade row compressor model study of a J85-13 engine operating with uniform and with circumferentially distorted inlet flow. Details of the geometry and the derived blade row characteristics used to simulate the clean inlet performance are given. A stability criterion based upon the self developing unsteady internal flows near surge provided an accurate determination of the clean inlet surge line. The basic model was modified to include an arbitrary extent multi-sector parallel compressor configuration for investigating 180 deg 1/rev total pressure, total temperature, and combined total pressure and total temperature distortions. The combined distortions included opposed, coincident, and 90 deg overlapped patterns. The predicted losses in surge pressure ratio matched the measured data trends at all speeds and gave accurate predictions at high corrected speeds where the slope of the speed lines approached the vertical.

Effect of spatial inlet temperature and pressure distortion on turbofan engine stability

NASA Technical Reports Server (NTRS)

Mehalic, Charles M.

1988-01-01

The effects of circumferential and radial inlet temperature distortion, circumferential pressure distortion, and combined temperature and pressure distortion on the stability of an advanced turbofan engine were investigated experimentally at simulated altitude conditions. With circumferential and radial inlet temperature distortion, a flow instability generated by the fan operating near stall caused the high-pressure compressor to surge at, or near, the same time as the fan. The effect of combined distortion was dependent on the relative location of the high-temperature and low-pressure regions; high-pressure compressor stalls occurred when the regions coincided, and fan stalls occurred with the regions separated.

Analysis of inlet flow distortion and turbulence effects on compressor stability

NASA Technical Reports Server (NTRS)

Melick, H. C., Jr.

1973-01-01

The effect of steady state circumferential total pressure distortion on the loss in compressor stall pressure ratio has been established by analytical techniques. Full scale engine and compressor/fan component test data were used to provide direct evaluation of the analysis. Specifically, since a circumferential total pressure distortion in an inlet system will result in unsteady flow in the coordinate system of the rotor blades, analysis of this type distortion must be performed from an unsteady aerodynamic point of view. By application of the fundamental aerothermodynamic laws to the inlet/compressor system, parameters important in the design of such a system for compatible operation have been identified. A time constant, directly related to the compressor rotor chord, was found to be significant, indicating compressor sensitivity to circumferential distortion is directly dependent on the rotor chord.

Two-stage fan. 3: Data and performance with rotor tip casing treatment, uniform and distorted inlet flows

NASA Technical Reports Server (NTRS)

Burger, G. D.; Hodges, T. R.; Keenan, M. J.

1975-01-01

A two stage fan with a 1st-stage rotor design tip speed of 1450 ft/sec, a design pressure ratio of 2.8, and corrected flow of 184.2 lbm/sec was tested with axial skewed slots in the casings over the tips of both rotors. The variable stagger stators were set in the nominal positions. Casing treatment improved stall margin by nine percentage points at 70 percent speed but decreased stall margin, efficiency, and flow by small amounts at design speed. Treatment improved first stage performance at low speed only and decreased second stage performance at all operating conditions. Casing treatment did not affect the stall line with tip radially distorted flow but improved stall margin with circumferentially distorted flow. Casing treatment increased the attenuation for both types of inlet flow distortion.

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.

Dynamics of flow control in an emulated boundary layer-ingesting offset diffuser

NASA Astrophysics Data System (ADS)

Gissen, A. N.; Vukasinovic, B.; Glezer, A.

2014-08-01

Dynamics of flow control comprised of arrays of active (synthetic jets) and passive (vanes) control elements , and its effectiveness for suppression of total-pressure distortion is investigated experimentally in an offset diffuser, in the absence of internal flow separation. The experiments are conducted in a wind tunnel inlet model at speeds up to M = 0.55 using approach flow conditioning that mimics boundary layer ingestion on a Blended-Wing-Body platform. Time-dependent distortion of the dynamic total-pressure field at the `engine face' is measured using an array of forty total-pressure probes, and the control-induced distortion changes are analyzed using triple decomposition and proper orthogonal decomposition (POD). These data indicate that an array of the flow control small-scale synthetic jet vortices merge into two large-scale, counter-rotating streamwise vortices that exert significant changes in the flow distortion. The two most energetic POD modes appear to govern the distortion dynamics in either active or hybrid flow control approaches. Finally, it is shown that the present control approach is sufficiently robust to reduce distortion with different inlet conditions of the baseline flow.

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 change of the inlet geometry of a centrifugal compressor stage and its influence on the compressor performance

NASA Astrophysics Data System (ADS)

Wang, Leilei; Yang, Ce; Zhao, Ben; Lao, Dazhong; Ma, Chaochen; Li, Du

2013-06-01

The impact on the compressor performance is important for designing the inlet pipe of the centrifugal compressor of a vehicle turbocharger with different inlet pipes. First, an experiment was performed to determine the compressor performance from three cases: a straight inlet pipe, a long bent inlet pipe and a short bent inlet pipe. Next, dynamic sensors were installed in key positions to collect the sign of the unsteady pressure of the centrifugal compressor. Combined with the results of numerical simulations, the total pressure distortion in the pipes, the pressure distributions on the blades and the pressure variability in the diffuser are studied in detail. The results can be summarized as follows: a bent pipe results in an inlet distortion to the compressor, which leads to performance degradation, and the effect is more apparent as the mass flow rate increases. The distortion induced by the bent inlet is not only influenced by the distance between the outlet of the bent section and the leading edge of the impeller but also by the impeller rotation. The flow fields in the centrifugal impeller and the diffuser are influenced by a coupling effect produced by the upstream inlet distortion and the downstream blocking effect from the volute tongue. If the inlet geometry is changed, the distributions and the fluctuation intensities of the static pressure on the main blade surface of the centrifugal impeller and in the diffuser are changed accordingly.

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.

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

NASA Technical Reports Server (NTRS)

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

2005-01-01

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

Comparison of Engine/Inlet Distortion Measurements with MEMS and ESP Pressure Sensors

NASA Technical Reports Server (NTRS)

Soto, Hector L.; Hernandez, Corey D.

2004-01-01

A study of active-flow control in a small-scale boundary layer ingestion inlet was conducted at the NASA Langley Basic Aerodynamic Research Tunnel (BART). Forty MEMS pressure sensors, in a rake style configuration, were used to examine both the mean (DC) and high frequency (AC) components of the total pressure across the inlet/engine interface plane. The mean component was acquired and used to calculate pressure distortion. The AC component was acquired separately, at a high sampling rate, and is used to study the unsteady effects of the active-flow control. An identical total pressure rake, utilizing an Electronically Scanned Pressure (ESP) system, was also used to calculate distortion; a comparison of the results obtained using the two rakes is presented.

Use of Vortex Generators to Reduce Distortion for Mach 1.6 Streamline-Traced Supersonic Inlets

NASA Technical Reports Server (NTRS)

Baydar, Ezgihan; Lu, Frank; Slater, John W.; Trefny, Chuck

2016-01-01

Reduce the total pressure distortion at the engine-fan face due to low-momentum flow caused by the interaction of an external terminal shock at the turbulent boundary layer along a streamline-traced external-compression (STEX) inlet for Mach 1.6.

Modeling and analysis of the TF30-P-3 compressor system with inlet pressure distortion

NASA Technical Reports Server (NTRS)

Mazzawy, R. S.; Banks, G. A.

1976-01-01

Circumferential inlet distortion testing of a TF30-P-3 afterburning turbofan engine was conducted at NASA-Lewis Research Center. Pratt and Whitney Aircraft analyzed the data using its multiple segment parallel compressor model and classical compressor theory. Distortion attenuation analysis resulted in a detailed flow field calculation with good agreement between multiple segment model predictions and the test data. Sensitivity of the engine stall line to circumferential inlet distortion was calculated on the basis of parallel compressor theory to be more severe than indicated by the data. However, the calculated stall site location was in agreement with high response instrumentation measurements.

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

Optimal Inlet Shape Design of N2B Hybrid Wing Body Configuration

NASA Technical Reports Server (NTRS)

Kim, Hyoungjin; Liou, Meng-Sing

2012-01-01

The N2B hybrid wing body aircraft was conceptually designed to meet environmental and performance goals for the N+2 generation transport set by the Subsonic Fixed Wing project of NASA Fundamental Aeronautics Program. In the present study, flow simulations are conducted around the N2B configuration by a Reynolds-averaged Navier-Stokes flow solver using unstructured meshes. Boundary conditions at engine fan face and nozzle exhaust planes are provided by the NPSS thermodynamic engine cycle model. The flow simulations reveal challenging design issues arising from boundary layer ingestion offset inlet and airframe-propulsion integration. Adjoint-based optimal designs are then conducted for the inlet shape to minimize the airframe drag force and flow distortion at fan faces. Design surfaces are parameterized by NURBS, and the cowl lip geometry is modified by a spring analogy approach. By the drag minimization design, flow separation on the cowl surfaces are almost removed, and shock wave strength got remarkably reduced. For the distortion minimization design, a circumferential distortion indicator DPCP(sub avg) is adopted as the design objective and diffuser bottom and side wall surfaces are perturbed for the design. The distortion minimization results in a 12.5 % reduction in the objective function.

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.

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.

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.

Effect on fan flow characteristics of length and axial location of a cascade thrust reverser

NASA Technical Reports Server (NTRS)

Dietrich, D. A.

1975-01-01

A series of static tests were conducted on a model fan with a diameter of 14.0 cm to determine the fan operating characteristics, the inlet static pressure contours, the fan-exit total and static pressure contours, and the fan-exit pressure distortion parameters associated with the installation of a partial-circumferential-emission cascade thrust reverser. The tests variables included the cascade axial length, the axial location of the reverser, and the type of fan inlet. It was shown that significant total and static pressure distortions were produced in the fan aft duct, and that some configurations induced a static pressure distortion at the fan face. The amount of flow passed by the fan and the level of the flow distortions were dependent upon all the variables tested.

Experimental Investigation of Inlet Distortion in a Multistage Axial Compressor

NASA Astrophysics Data System (ADS)

Rusu, Razvan

The primary objective of this research is to present results and methodologies used to study total pressure inlet distortion in a multi-stage axial compressor environment. The study was performed at the Purdue 3-Stage Axial Compressor Facility (P3S) which models the final three stages of a production turbofan engine's high-pressure compressor (HPC). The goal of this study was twofold; first, to design, implement, and validate a circumferentially traversable total pressure inlet distortion generation system, and second, to demonstrate data acquisition methods to characterize the inter-stage total pressure flow fields to study the propagation and attenuation of a one-per-rev total pressure distortion. The datasets acquired for this study are intended to support the development and validation of novel computational tools and flow physics models for turbomachinery flow analysis. Total pressure inlet distortion was generated using a series of low-porosity wire gauze screens placed upstream of the compressor in the inlet duct. The screens are mounted to a rotatable duct section that can be precisely controlled. The P3S compressor features fixed instrumentation stations located at the aerodynamic interface plane (AIP) and downstream and upstream of each vane row. Furthermore, the compressor features individually indexable stator vanes which can be traverse by up to two vane passages. Using a series of coordinated distortion and vane traverses, the total pressure flow field at the AIP and subsequent inter-stage stations was characterized with a high circumferential resolution. The uniformity of the honeycomb carrier was demonstrated by characterizing the flow field at the AIP while no distortion screens where installed. Next, the distortion screen used for this study was selected following three iterations of porosity reduction. The selected screen consisted of a series of layered screens with a 100% radial extent and a 120° circumferential extent. A detailed total pressure flow field characterization of the AIP was performed using the selected screen at nominal, low, and high compressor loading. Thermal anemometry was used to characterize the spatial variation in turbulence intensity at the AIP in an effort to further define inlet boundary conditions for future computational investigations. Two data acquisition methods for the study of distortion propagation and attenuation were utilized in this study. The first method approximated the bulk flow through each vane passage using a single rake measurement positioned near the center of the passage. All vane passages were measured virtually by rotating the distortion upstream by an increment equal to one vane passage. This method proved successful in tracking the distortion propagation and attenuation from the AIP up until the compressor exit. A second, more detailed, inter-stage flow field characterization method was used that generated a total pressure field with a circumferential resolution of 880 increments, or one every 0.41°. The resulting fields demonstrated the importance of secondary flows in the propagation of a total pressure distortion at the different loading conditions investigated. A second objective of this research was to document proposals and design efforts to outfit the existing P3S research compressor with a strain gage telemetry system. The purpose of this system is to validate and supplement existing blade tip timing data on the embedded rotor stage to support the development and validation of novel aeromechanical analysis tools. Integration strategies and telemetry considerations are discussed based on proposals and consultation provided by suppliers.

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.

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

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

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

2000-01-01

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

Control of Inflow Distortion in a Scarf Inlet

NASA Technical Reports Server (NTRS)

Gerhold, Carl H.; Clark, Lorenzo R.; Biedron, Robert T.

2002-01-01

The scarf inlet has the potential to reduce aircraft inlet noise radiation to the ground by reflecting it into the space above the engine. Without forward motion of the engine, the non-symmetry of the inlet causes inflow distortion which generates noise that is greater than the noise reduction of the scarf. However, acoustic evaluations of aircraft engines are often done on static test stands. A method to reduce inflow distortion by boundary layer suction is proposed and evaluated using a model of a high bypass ratio engine located in an anechoic chamber. The design goal of the flow control system is to make the inflow to the inlet circumferentially uniform and to eliminate reversed flow. This minimizes the inflow distortion and allows for acoustic evaluation of the scarf inlet on a static test stand. The inlet boundary layer suction effectiveness is evaluated both by aerodynamic and by acoustic measurements. Although the design goal is not met, the control system is found to have a beneficial effect on the engine operation, reducing blade stall and speed variation. This is quantified by two acoustic benefits, reduction both of the variability of tone noise and of the low frequency wideband noise due to the inflow distortion. It is felt that a compromise in the manufacture of the control hardware contributes to the inability of the control system to perform as expected from the analysis. The control system with sufficient authority is felt to have the potential to permit reliable acoustic testing in a static configuration of engines with non-symmetric inlets. Because the control system can improve operation of the engine, it may also have the potential to reduce noise and vibration and enhance engine longevity during low speed ground operations in the terminal area.

Application of Computational Fluid Dynamics to the Study of Vortex Flow Control for the Management of Inlet Distortion

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Gibb, James

1992-01-01

The present study demonstrates that the Reduced Navier-Stokes code RNS3D can be used very effectively to develop a vortex generator installation for the purpose of minimizing the engine face circumferential distortion by controlling the development of secondary flow. The computing times required are small enough that studies such as this are feasible within an analysis-design environment with all its constraints of time and costs. This research study also established the nature of the performance improvements that can be realized with vortex flow control, and suggests a set of aerodynamic properties (called observations) that can be used to arrive at a successful vortex generator installation design. The ultimate aim of this research is to manage inlet distortion by controlling secondary flow through an arrangements of vortex generators configurations tailored to the specific aerodynamic characteristics of the inlet duct. This study also indicated that scaling between flight and typical wind tunnel test conditions is possible only within a very narrow range of generator configurations close to an optimum installation. This paper also suggests a possible law that can be used to scale generator blade height for experimental testing, but further research in this area is needed before it can be effectively applied to practical problems. Lastly, this study indicated that vortex generator installation design for inlet ducts is more complex than simply satisfying the requirement of attached flow, it must satisfy the requirement of minimum engine face distortion.

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.

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.

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.

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.

Cyclic Symmetry Finite Element Forced Response Analysis of a Distortion-Tolerant Fan with Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Min, J. B.; Reddy, T. S. R.; Bakhle, M. A.; Coroneos, R. M.; Stefko, G. L.; Provenza, A. J.; Duffy, K. P.

2018-01-01

Accurate prediction of the blade vibration stress is required to determine overall durability of fan blade design under Boundary Layer Ingestion (BLI) distorted flow environments. Traditional single blade modeling technique is incapable of representing accurate modeling for the entire rotor blade system subject to complex dynamic loading behaviors and vibrations in distorted flow conditions. A particular objective of our work was to develop a high-fidelity full-rotor aeromechanics analysis capability for a system subjected to a distorted inlet flow by applying cyclic symmetry finite element modeling methodology. This reduction modeling method allows computationally very efficient analysis using a small periodic section of the full rotor blade system. Experimental testing by the use of the 8-foot by 6-foot Supersonic Wind Tunnel Test facility at NASA Glenn Research Center was also carried out for the system designated as the Boundary Layer Ingesting Inlet/Distortion-Tolerant Fan (BLI2DTF) technology development. The results obtained from the present numerical modeling technique were evaluated with those of the wind tunnel experimental test, toward establishing a computationally efficient aeromechanics analysis modeling tool facilitating for analyses of the full rotor blade systems subjected to a distorted inlet flow conditions. Fairly good correlations were achieved hence our computational modeling techniques were fully demonstrated. The analysis result showed that the safety margin requirement set in the BLI2DTF fan blade design provided a sufficient margin with respect to the operating speed range.

Performance of a 1.57 pressure-ratio transonic fan stage with a screen-induced 90 deg circumferential inlet flow distortion

NASA Technical Reports Server (NTRS)

Sanger, N. L.

1976-01-01

A transonic fan stage having a design pressure ratio of 1.57 was tested with a 90 degree circumferential distortion imposed on the inlet flow. The rotor diameter was approximately 50.8 cm, and the design pressure ratio was 1.60 at a tip speed of 425 m/sec. Overall performance at 70 and 100 percent of design speed showed a loss of stall pressure ratio and flow range at design speed and no significant loss in stall pressure ratio at 70 percent of design speed. Detailed flow measurements are presented to show the rotor-upstream flow interactions and the attenuation and amplification properties through the stage.

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.

Preliminary Data on the Effects of Inlet Pressure Distortions on the J57-P-1 Turbojet Engine

NASA Technical Reports Server (NTRS)

Wallner, Lewis E.; Lubick, Robert J.; Einstein, Thomas H.

1954-01-01

An investigation to determine the steady-state and surge characteristics of the J57-P-1 two-spool turbojet engine with various inlet air-flow distortions was conducted in the altitude wind tunnel at the NACA Lewis laboratory. Along with a uniform inlet total-pressure distribution, one circumferential and three radial pressure distortions were investigated. Data were obtained over a complete range of compressor speeds both with and without intercompressor air bleed at a flight Mach number of 0.8 and at altitudes of 35,000 and 50,000 feet. Total-pressure distortions of the magnitudes investigated had very little effect on the steady-state operating line for either the outer or inner compressor. The small radial distortions investigated also had engine over that obtained with the uniform inlet pressure distribution. The circumferential distortion, however, raised the minimum speed at which the engine could operate without encountering surge when the intercompressor bleeds were closed. This increase in minimum speed resulted in a substantial reduction in the operable speed range accompanied by a reduction in the altitude operating limit.

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.

Optimization of pressure probe placement and data analysis of engine-inlet distortion

NASA Astrophysics Data System (ADS)

Walter, S. F.

The purpose of this research is to examine methods by which quantification of inlet flow distortion may be improved upon. Specifically, this research investigates how data interpolation effects results, optimizing sampling locations of the flow, and determining the sensitivity related to how many sample locations there are. The main parameters that are indicative of a "good" design are total pressure recovery, mass flow capture, and distortion. This work focuses on the total pressure distortion, which describes the amount of non-uniformity that exists in the flow as it enters the engine. All engines must tolerate some level of distortion, however too much distortion can cause the engine to stall or the inlet to unstart. Flow distortion is measured at the interface between the inlet and the engine. To determine inlet flow distortion, a combination of computational and experimental pressure data is generated and then collapsed into an index that indicates the amount of distortion. Computational simulations generate continuous contour maps, but experimental data is discrete. Researchers require continuous contour maps to evaluate the overall distortion pattern. There is no guidance on how to best manipulate discrete points into a continuous pattern. Using one experimental, 320 probe data set and one, 320 point computational data set with three test runs each, this work compares the pressure results obtained using all 320 points of data from the original sets, both quantitatively and qualitatively, with results derived from selecting 40 grid point subsets and interpolating to 320 grid points. Each of the two, 40 point sets were interpolated to 320 grid points using four different interpolation methods in an attempt to establish the best method for interpolating small sets of data into an accurate, continuous contour map. Interpolation methods investigated are bilinear, spline, and Kriging in Cartesian space, as well as angular in polar space. Spline interpolation methods should be used as they result in the most accurate, precise, and visually correct predictions when compared results achieved from the full data sets. Researchers were interested if fewer than the recommended 40 probes could be used - especially when placed in areas of high interest - but still obtain equivalent or better results. For this investigation, the computational results from a two-dimensional inlet and experimental results of an axisymmetric inlet were used. To find the areas of interest, a uniform sampling of all possible locations was run through a Monte Carlo simulation with a varying number of probes. A probability density function of the resultant distortion index was plotted. Certain probes are required to come within the desired accuracy level of the distortion index based on the full data set. For the experimental results, all three test cases could be characterized with 20 probes. For the axisymmetric inlet, placing 40 probes in select locations could get the results for parameters of interest within less than 10% of the exact solution for almost all cases. For the two dimensional inlet, the results were not as clear. 80 probes were required to get within 10% of the exact solution for all run numbers, although this is largely due to the small value of the exact result. The sensitivity of each probe added to the experiment was analyzed. Instead of looking at the overall pattern established by optimizing probe placements, the focus is on varying the number of sampled probes from 20 to 40. The number of points falling within a 1% tolerance band of the exact solution were counted as good points. The results were normalized for each data set and a general sensitivity function was found to determine the sensitivity of the results. A linear regression was used to generalize the results for all data sets used in this work. However, they can be used by directly comparing the number of good points obtained with various numbers of probes as well. The sensitivity in the results is higher when fewer probes are used and gradually tapers off near 40 probes. There is a bigger gain in good points when the number of probes is increased from 20 to 21 probes than from 39 to 40 probes.

Vortex generator design for aircraft inlet distortion as a numerical optimization problem

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Levy, Ralph

1991-01-01

Aerodynamic compatibility of aircraft/inlet/engine systems is a difficult design problem for aircraft that must operate in many different flight regimes. Takeoff, subsonic cruise, supersonic cruise, transonic maneuvering, and high altitude loiter each place different constraints on inlet design. Vortex generators, small wing like sections mounted on the inside surfaces of the inlet duct, are used to control flow separation and engine face distortion. The design of vortex generator installations in an inlet is defined as a problem addressable by numerical optimization techniques. A performance parameter is suggested to account for both inlet distortion and total pressure loss at a series of design flight conditions. The resulting optimization problem is difficult since some of the design parameters take on integer values. If numerical procedures could be used to reduce multimillion dollar development test programs to a small set of verification tests, numerical optimization could have a significant impact on both cost and elapsed time to design new aircraft.

Vortex Generators in a Two-Dimensional, External-Compression Supersonic Inlet

NASA Technical Reports Server (NTRS)

Baydar, Ezgihan; Lu, Frank K.; Slater, John W.

2016-01-01

Vortex generators within a two-dimensional, external-compression supersonic inlet for Mach 1.6 were investigated to determine their ability to increase total pressure recovery, reduce total pressure distortion, and improve the boundary layer. The vortex generators studied included vanes and ramps. The geometric factors of the vortex generators studied included height, length, spacing, and positions upstream and downstream of the inlet terminal shock. The flow through the inlet was simulated through the computational solution of the steady-state Reynolds-averaged Navier-Stokes equations on multi-block, structured grids. The vortex generators were simulated by either gridding the geometry of the vortex generators or modeling the vortices generated by the vortex generators. The inlet performance was characterized by the inlet total pressure recovery, total pressure distortion, and incompressible shape factor of the boundary-layer at the engine face. The results suggested that downstream vanes reduced the distortion and improved the boundary layer. The height of the vortex generators had the greatest effect of the geometric factors.

Isolated rotor noise due to inlet distortion or turbulence

NASA Technical Reports Server (NTRS)

Mani, R.

1974-01-01

Theoretical formulation, analysis, and results are presented that are necessary to analyze quadrupole noise generated from a loaded, subsonic rotor because of its interaction with an inflow distortion or inlet turbulence. The ratio of quadrupole to dipole noise is largely a function of the axial Mach number, wheel tip Mach number, rotor solidity, and total pressure ratio across the rotor. It is relatively independent of the specific form of the inflow distortion or inlet turbulence. Comparisons with experimental data only succeed in predicting gross levels at a given speed and fail to predict the variation of noise at fixed speed with flow and pressure ratio. Likely sources of this discrepancy are discussed.

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.

Compressor Performance Scaling in the Presence of Non-Uniform Flow

NASA Astrophysics Data System (ADS)

Hill, David Jarrod

Fuselage-embedded engines in future aircraft will see increased flow distortions due to the ingestion of airframe boundary layers. This reduces the required propulsive power compared to podded engines. Inlet flow distortions mean that localized regions of flow within the fan and first stage compressor are operating at off-design conditions. It is important to weigh the benefit of increased vehicle propulsive efficiency against the resultant reduction in engine efficiency. High computational cost has limited most past research to single distortion studies. The objective of this thesis is to extract scaling laws for transonic compressor performance in the presence of various distortion patterns and intensities. The machine studied is the NASA R67 transonic compressor. Volumetric source terms are used to model rotor and stator blade rows. The modelling approach is an innovative combination of existing flow turning and loss models, combined with a compressible flow correction. This approach allows for a steady calculation to capture distortion transfer; as a result, the computational cost is reduced by two orders of magnitude. At peak efficiency, the rotor work coefficient and isentropic efficiency are matched within 1.4% of previously published experimental results. A key finding of this thesis is that, in non-uniform flow, the state-of-the-art loss model employed is unable to capture the impact of variations in local flow coefficient, limiting the analysis of local entropy generation. New insight explains the mechanism governing the interaction between a total temperature distortion and a compressor rotor. A parametric study comprising 16 inlet distortions reveals that for total temperature distortions, upstream flow redistribution and rotor diffusion factor changes are shown to scale linearly with distortion severity. Linear diffusion factor scaling does not hold true for total pressure distortions. For combined total temperature and total pressure distortions, the changes in rotor diffusion factor are predicted by the summation of the individual distortions, within 3.65%.

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.

S-Duct Engine Inlet Flow Control Using SDBD Plasma Streamwise Vortex Generators

NASA Astrophysics Data System (ADS)

Kelley, Christopher; He, Chuan; Corke, Thomas

2009-11-01

The results of a numerical simulation and experiment characterizing the performance of plasma streamwise vortex generators in controlling separation and secondary flow within a serpentine, diffusing duct are presented. A no flow control case is first run to check agreement of location of separation, development of secondary flow, and total pressure recovery between the experiment and numerical results. Upon validation, passive vane-type vortex generators and plasma streamwise vortex generators are implemented to increase total pressure recovery and reduce flow distortion at the aerodynamic interface plane: the exit of the S-duct. Total pressure recovery is found experimentally with a pitot probe rake assembly at the aerodynamic interface plane. Stagnation pressure distortion descriptors are also presented to show the performance increase with plasma streamwise vortex generators in comparison to the baseline no flow control case. These performance parameters show that streamwise plasma vortex generators are an effective alternative to vane-type vortex generators in total pressure recovery and total pressure distortion reduction in S-duct inlets.

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.

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

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

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.

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.

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.

Coupled Control of Flow Separation and Streamwise Vortical Structures

NASA Astrophysics Data System (ADS)

Burrows, Travis; Vukasinovic, Bojan; Glezer, Ari

2017-11-01

The flow in offset diffusers of modern propulsion systems are dominated by streamwise vorticity concentrations that advect of low-momentum fluid from the flow boundaries into the core flow and give rise to flow distortion and losses at the engine inlet. Because the formation of these vortices is strongly coupled to trapped vorticity concentrations within locally-separated flow domains over concave surfaces of the diffuser bends, this coupling is exploited for controlling the streamwise evolution of the vortices and thereby significantly reduce the flow distortion and losses. The scale and topology of the trapped vorticity are manipulated at an operating throat Mach number of 0.64 by using a spanwise array of fluidic oscillating jets that are placed upstream of the separation domain. The present investigations demonstrate that the actuation alters the structure of both the trapped and streamwise vortices. In particular, the distribution of the streamwise vortices is altered and their strength is diminished by actuation-induced streamwise vorticity concentrations of opposite sense. As a result, the actuation leads to significant suppression of pressure distortion at the engine inlet (by as much as 60%) at an actuation level that utilizes less than 0.4% of the diffuser's mass flow rate. Supported by ONR.

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.

Effect of Blowing on Boundary Layer of Scarf Inlet

NASA Technical Reports Server (NTRS)

Gerhold, Carl H.; Clark, Lorenzo R.

2004-01-01

When aircraft operate in stationary or low speed conditions, airflow into the engine accelerates around the inlet lip and pockets of turbulence that cause noise and vibration can be ingested. This problem has been encountered with engines equipped with the scarf inlet, both in full scale and in model tests, where the noise produced during the static test makes it difficult to assess the noise reduction performance of the scarf inlet. NASA Langley researchers have implemented boundary layer control in an attempt to reduce the influence of the flow nonuniformity in a 12-in. diameter model of a high bypass fan engine mounted in an anechoic chamber. Static pressures and boundary layer profiles were measured in the inlet and far field acoustic measurements were made to assess the effectiveness of the blowing treatment. The blowing system was found to lack the authority to overcome the inlet distortions. Methods to improve the implementation of boundary layer control to reduce inlet distortion are discussed.

A comparative study of full Navier-Stokes and Reduced Navier-Stokes analyses for separating flows within a diffusing inlet S-duct

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Reddy, D. R.; Kapoor, K.

1993-01-01

A three-dimensional implicit Full Navier-Stokes (FNS) analysis and a 3D Reduced Navier-Stokes (RNS) initial value space marching solution technique has been applied to a class of separate flow problems within a diffusing S-duct configuration characterized as vortex-liftoff. Both Full Navier-Stokes and Reduced Navier-Stokes solution techniques were able to capture the overall flow physics of vortex lift-off, however more consideration must be given to the development of turbulence models for the prediction of the locations of separation and reattachment. This accounts for some of the discrepancies in the prediction of the relevant inlet distortion descriptors, particularly circumferential distortion. The 3D RNS solution technique adequately described the topological structure of flow separation associated with vortex lift-off.

Unsteady Reynolds-averaged Navier-Stokes simulations of inlet distortion in the fan system of a gas-turbine aero-engine

NASA Astrophysics Data System (ADS)

Spotts, Nathan

As modern trends in commercial aircraft design move toward high-bypass-ratio fan systems of increasing diameter with shorter, nonaxisymmetric nacelle geometries, inlet distortion is becoming common in all operating regimes. The distortion may induce aerodynamic instabilities within the fan system, leading to catastrophic damage to fan blades, should the surge margin be exceeded. Even in the absence of system instability, the heterogeneity of the flow affects aerodynamic performance significantly. Therefore, an understanding of fan-distortion interaction is critical to aircraft engine system design. This thesis research elucidates the complex fluid dynamics and fan-distortion interaction by means of computational fluid dynamics (CFD) modeling of a complete engine fan system; including rotor, stator, spinner, nacelle and nozzle; under conditions typical of those encountered by commercial aircraft. The CFD simulations, based on a Reynolds-averaged Navier-Stokes (RANS) approach, were unsteady, three-dimensional, and of a full-annulus geometry. A thorough, systematic validation has been performed for configurations from a single passage of a rotor to a full-annulus system by comparing the predicted flow characteristics and aerodynamic performance to those found in literature. The original contributions of this research include the integration of a complete engine fan system, based on the NASA rotor 67 transonic stage and representative of the propulsion systems in commercial aircraft, and a benchmark case for unsteady RANS simulations of distorted flow in such a geometry under realistic operating conditions. This study is unique in that the complex flow dynamics, resulting from fan-distortion interaction, were illustrated in a practical geometry under realistic operating conditions. For example, the compressive stage is shown to influence upstream static pressure distributions and thus suppress separation of flow on the nacelle. Knowledge of such flow physics is valuable for engine system design.

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 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.

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.

Distribution analysis for F100(3) engine

NASA Technical Reports Server (NTRS)

Walter, W. A.; Shaw, M.

1980-01-01

The F100(3) compression system response to inlet circumferential distortion was investigated using an analytical compressor flow model. Compression system response to several types of distortion, including pressure, temperature, and combined pressure/temperature distortions, was investigated. The predicted response trends were used in planning future F100(3) distortion tests. Results show that compression system response to combined temperature and pressure distortions depends upon the relative orientation, as well as the individual amplitudes and circumferential extents of the distortions. Also the usefulness of the analytical predictions in planning engine distortion tests is indicated.

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.

Performance with and without inlet radial distortion of a transonic fan stage designed for reduced loading in the tip region

NASA Technical Reports Server (NTRS)

Schmidt, J. F.; Ruggeri, R. S.

1978-01-01

A transonic compressor stage designed for a reduced loading in the tip region of the rotor blades was tested with and without inlet radial distortion. The rotor was 50 cm in diameter and designed for an operating tip speed of 420 m/sec. Although the rotor blade loading in the tip region was reduced to provide additional operating range, analysis of the data indicates that the flow around the damper appears to be critical and limited the stable operating range of this stage. For all levels of tip and hub radial distortion, there was a large reduction in the rotor stall margin.

High-loading low-speed fan study. 4: Data and performance with redesign stator and including a rotor tip casing treatment

NASA Technical Reports Server (NTRS)

Harley, K. G.; Odegard, P. A.; Burdsall, E. A.

1972-01-01

A single stage fan with a rotor tip speed of 1000 ft/sec(304.8 m/sec) and a hub-to-tip ratio of 0.392 was retested with a redesigned stator. Tests were conducted with uniform inlet, tip-radial, hub-radial, and circumferential inlet distortions. With uniform inlet flow, stall margin was improved 12 percentage points above that with the original stator. The fan demonstrated an efficiency of 0.883 and a stall margin of 15 percent at a pressure ratio of 1.488 and a specific flow of 41.17 lb/sec/sq ft. Tests were also made with a redesigned casing treatment consisting of skewed slots over the rotor blade tips. This casing treatment gave a 7 percentage point improvement in stall margin when tested with tip radial distortion (when the rotor tip initiated stall). Noise measurements at the fan inlet and exit indicate no effect from closing the stator 10 degrees, nor were there measurable effects from adding skewed slots over the blade tips.

Inlet Distortion Generation for a Transonic Compressor

DTIC Science & Technology

2004-09-01

9 Figure 6. Compressor pumping characteristic measured at 90% design speed and degradation assumed for distortion design ...INTENTIONALLY LEFT BLANK 1 I. INTRODUCTION Engines for military fighter aircraft must be designed to operate stably over a required flight envelope. An...adequate “stall margin” is usually an engine design requirement. Since distortion of the flow into the fan or compressor is known to reduce the

Analysis of internal flow of J85-13 multistage compressor

NASA Technical Reports Server (NTRS)

Hager, R. D.

1977-01-01

Interstage data recorded on a J85-13 engine were used to analyze the internal flow of the compressor. Measured pressures and temperatures were used as input to a streamline analysis program to calculate the velocity diagrams at the inlet and outlet of each blade row. From the velocity diagrams and blade geometry, selected blade-element performance parameters were calculated. From the detailed analysis it is concluded that the compressor is probably hub critical (stall initiates at the hub) in the latter stages for the design speed conditions. As a result, the casing treatment over the blade tips has little or no effect on stall margin at design speed. Radial inlet distortion did not appear to change the flow in the stages that control stall because of the rapid attenuation of the distortion within the compressor.

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

NASA Technical Reports Server (NTRS)

Woollett, R. R.

1983-01-01

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

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

NASA Technical Reports Server (NTRS)

Woollett, R. R.

1983-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

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 investigation of bleed configurations and their effect on shock wave/boundary layer interactions

NASA Technical Reports Server (NTRS)

Hamed, Awatef

1995-01-01

The design of high efficiency supersonic inlets is a complex task involving the optimization of a number of performance parameters such as pressure recovery, spillage, drag, and exit distortion profile, over the flight Mach number range. Computational techniques must be capable of accurately simulating the physics of shock/boundary layer interactions, secondary corner flows, flow separation, and bleed if they are to be useful in the design. In particular, bleed and flow separation, play an important role in inlet unstart, and the associated pressure oscillations. Numerical simulations were conducted to investigate some of the basic physical phenomena associated with bleed in oblique shock wave boundary layer interactions that affect the inlet performance.

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.

A design strategy for the use of vortex generators to manage inlet-engine distortion using computational fluid dynamics

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Levy, Ralph

1991-01-01

A reduced Navier-Stokes solution technique was successfully used to design vortex generator installations for the purpose of minimizing engine face distortion by restructuring the development of secondary flow that is induced in typical 3-D curved inlet ducts. The results indicate that there exists an optimum axial location for this installation of corotating vortex generators, and within this configuration, there exists a maximum spacing between generator blades above which the engine face distortion increases rapidly. Installed vortex generator performance, as measured by engine face circumferential distortion descriptors, is sensitive to Reynolds number and thereby the generator scale, i.e., the ratio of generator blade height to local boundary layer thickness. Installations of corotating vortex generators work well in terms of minimizing engine face distortion within a limited range of generator scales. Hence, the design of vortex generator installations is a point design, and all other conditions are off design. In general, the loss levels associated with a properly designed vortex generator installation are very small; thus, they represent a very good method to manage engine face distortion. This study also showed that the vortex strength, generator scale, and secondary flow field structure have a complicated and interrelated influence over engine face distortion, over and above the influence of the initial arrangement of generators.

Mitigation of Adverse Effects Caused by Shock Wave Boundary Layer Interactions Through Optimal Wall Shaping

NASA Technical Reports Server (NTRS)

Liou, May-Fun; Lee, Byung Joon

2013-01-01

It is known that the adverse effects of shock wave boundary layer interactions in high speed inlets include reduced total pressure recovery and highly distorted flow at the aerodynamic interface plane (AIP). This paper presents a design method for flow control which creates perturbations in geometry. These perturbations are tailored to change the flow structures in order to minimize shock wave boundary layer interactions (SWBLI) inside supersonic inlets. Optimizing the shape of two dimensional micro-size bumps is shown to be a very effective flow control method for two-dimensional SWBLI. In investigating the three dimensional SWBLI, a square duct is employed as a baseline. To investigate the mechanism whereby the geometric elements of the baseline, i.e. the bottom wall, the sidewall and the corner, exert influence on the flow's aerodynamic characteristics, each element is studied and optimized separately. It is found that arrays of micro-size bumps on the bottom wall of the duct have little effect in improving total pressure recovery though they are useful in suppressing the incipient separation in three-dimensional problems. Shaping sidewall geometry is effective in re-distributing flow on the side wall and results in a less distorted flow at the exit. Subsequently, a near 50% reduction in distortion is achieved. A simple change in corner geometry resulted in a 2.4% improvement in total pressure recovery.

Characterization of Aerodynamic Performance of Boundary-Layer-Ingesting Inlet Under Crosswind

NASA Technical Reports Server (NTRS)

Liou, Meng-Sing; Lee, Byung Joon

2012-01-01

NASA has been studying future transport concepts, envisioned to be technically realizable in the timeframe of 2020-2030, to meet environmental and performance goals. One concept receiving considerable interest involves a propulsion system embedded into a hybrid wing-body aircraft. While offering significant advantages in fuel savings and noise reduction by this concept, there are several technical challenges that are not encountered in the current fleet and must be overcome so as to deliver target performance and operability. One of these challenges is associated with an inlet system that ingests a significantly thick boundary layer, developing along the wing-body surface, into a serpentine diffuser before the flow meeting fan blades. The flow is subject to considerable total pressure loss and distorted at the fan face, much more significantly than in the inlet system of conventional aircraft. In our previous studies [1, 2], we have shown that through innovative design changes on the airframe surface, it is possible to simultaneously increase total pressure recovery and decrease distortion in the flow, without resorting to conventional penalty-ridden flow control concepts, such as vortex generator or boundary layer bleeding/suction. In the current study, we are interested in understanding the following issues: how the embedded propulsion system performs under a crosswind condition by studying in detail the flow characteristics of two inlets, the baseline and another optimized previously under the cruise condition. With the insight, it is hoped that it can help in the follow-on study by devising effective strategies to minimize flow distortion arising from the integration of an embedded-engine system into an airframe to the level acceptable to the operation and fuel consumption before 2030. To achieve these demanding goals, non-conventional concepts are called for; but technology gap is too big that it requires evolutionary approach by focusing various concepts and technologies needed in the next three generations of aircraft, respectively named as N+1, N+2, and N+3. Noticeably, considerable reduction in each category of 1 is required in N+2 (relative to Boeing 777-200 and GE90 engines) and N+3 (relative to Boeing 737-800 and CFM56-7B engines). In this study, concepts for N+2 is our interest. A concept that has potential to achieve these metrics and has been under intensive study is the hybrid wing body (HWB) airframe with a tightly integrated propulsion system, see Fig. 1. The inlet is non-circular at the entrance and the entering flow, no longer uniform or free of disturbances, and is now carrying with it a boundary layer developing along the fuselage; the inlet is thus known as boundary-layer-ingesting (BLI) inlet.

A Full Navier-Stokes Analysis of Subsonic Diffuser of a Bifurcated 70/30 Supersonic Inlet for High Speed Civil Transport Application

NASA Technical Reports Server (NTRS)

Kapoor, Kamlesh; Anderson, Bernhard H.; Shaw, Robert J.

1994-01-01

A full Navier-Stokes analysis was performed to evaluate the performance of the subsonic diffuser of a NASA Lewis Research Center 70/30 mixed-compression bifurcated supersonic inlet for high speed civil transport application. The PARC3D code was used in the present study. The computations were also performed when approximately 2.5 percent of the engine mass flow was allowed to bypass through the engine bypass doors. The computational results were compared with the available experimental data which consisted of detailed Mach number and total pressure distribution along the entire length of the subsonic diffuser. The total pressure recovery, flow distortion, and crossflow velocity at the engine face were also calculated. The computed surface ramp and cowl pressure distributions were compared with experiments. Overall, the computational results compared well with experimental data. The present CFD analysis demonstrated that the bypass flow improves the total pressure recovery and lessens flow distortions at the engine face.

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.

Three-dimensional Aerodynamic Instability in Multi-stage Axial Compressors

NASA Technical Reports Server (NTRS)

Suder, Kenneth (Technical Monitor); Tan, Choon-Sooi

2003-01-01

Four separate tasks are reported. The first task: A Computational Model for Short Wavelength Stall Inception and Development In Multi-Stage Compressors; the second task: Three-dimensional Rotating Stall Inception and Effects of Rotating Tip Clearance Asymmetry in Axial Compressors; the third task:Development of an Effective Computational Methodology for Body Force Representation of High-speed Rotor 37; and the fourth task:Development of Circumferential Inlet Distortion through a Representative Eleven Stage High-speed axial compressor. The common theme that threaded throughout these four tasks is the conceptual framework that consists of quantifying flow processes at the fadcompressor blade passage level to define the compressor performance characteristics needed for addressing physical phenomena such compressor aerodynamic instability and compressor response to flow distoriton with length scales larger than compressor blade-to-blade spacing at the system level. The results from these two levels can be synthesized to: (1) simulate compressor aerodynamic instability inception local to a blade rotor tip and its development from a local flow event into the nonlinear limit cycle instability that involves the entire compressor as was demonstrated in the first task; (2) determine the conditions under which compressor stability assessment based on two-dimensional model may not be adequate and the effects of self-induced flow distortion on compressor stability limit as in the second task; (3) quantify multistage compressor response to inlet distortion in stagnation pressure as illustrated in the fourth task; and (4) elucidate its potential applicability for compressor map generation under uniform as well as non-uniform inlet flow given three-dimensional Navier-Stokes solution for each individual blade row as was demonstrated in the third task.

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

NASA Technical Reports Server (NTRS)

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

2003-01-01

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

Numerical Prediction of the Influence of Thrust Reverser on Aeroengine's Aerodynamic Stability

NASA Astrophysics Data System (ADS)

Zhiqiang, Wang; Xigang, Shen; Jun, Hu; Xiang, Gao; Liping, Liu

2017-11-01

A numerical method was developed to predict the aerodynamic stability of a high bypass ratio turbofan engine, at the landing stage of a large transport aircraft, when the thrust reverser was deployed. 3D CFD simulation and 2D aeroengine aerodynamic stability analysis code were performed in this work, the former is to achieve distortion coefficient for the analysis of engine stability. The 3D CFD simulation was divided into two steps, the single engine calculation and the integrated aircraft and engine calculation. Results of the CFD simulation show that with the decreasing of relative wind Mach number, the engine inlet will suffer more severe flow distortion. The total pressure and total temperature distortion coefficients at the inlet of the engines were obtained from the results of the numerical simulation. Then an aeroengine aerodynamic stability analysis program was used to quantitatively analyze the aerodynamic stability of the high bypass ratio turbofan engine. The results of the stability analysis show that the engine can work stably, when the reverser flow is re-ingested. But the anti-distortion ability of the booster is weaker than that of the fan and high pressure compressor. It is a weak link of engine stability.

Effects of inlet distortion on gas turbine combustion chamber exit temperature profiles

NASA Astrophysics Data System (ADS)

Maqsood, Omar Shahzada

Damage to a nozzle guide vane or blade, caused by non-uniform temperature distributions at the combustion chamber exit, is deleterious to turbine performance and can lead to expensive and time consuming overhaul and repair. A test rig was designed and constructed for the Allison 250-C20B combustion chamber to investigate the effects of inlet air distortion on the combustion chamber's exit temperature fields. The rig made use of the engine's diffuser tubes, combustion case, combustion liner, and first stage nozzle guide vane shield. Rig operating conditions simulated engine cruise conditions, matching the quasi-non-dimensional Mach number, equivalence ratio and Sauter mean diameter. The combustion chamber was tested with an even distribution of inlet air and a 4% difference in airflow at either side. An even distribution of inlet air to the combustion chamber did not create a uniform temperature profile and varying the inlet distribution of air exacerbated the profile's non-uniformity. The design of the combustion liner promoted the formation of an oval-shaped toroidal vortex inside the chamber, creating localized hot and cool sections separated by 90° that appeared in the exhaust. Uneven inlet air distributions skewed the oval vortex, increasing the temperature of the hot section nearest the side with the most mass flow rate and decreasing the temperature of the hot section on the opposite side. Keywords: Allison 250, Combustion, Dual-Entry, Exit Temperature Profile, Gas Turbine, Pattern Factor, Reverse Flow.

Circumferential distortion modeling of the TF30-P-3 compression system

NASA Technical Reports Server (NTRS)

Mazzawy, R. S.; Banks, G. A.

1977-01-01

Circumferential inlet pressure and temperature distortion testing of the TF30 P-3 turbofan engine was conducted. The compressor system at the test conditions run was modelled according to a multiple segment parallel compressor model. Aspects of engine operation and distortion configuration modelled include the effects of compressor bleeds, relative pressure-temperature distortion alignment and circumferential distortion extent. Model predictions for limiting distortion amplitudes and flow distributions within the compression system were compared with test results in order to evaluate predicted trends. Relatively good agreement was obtained. The model also identified the low pressure compressor as the stall-initiating component, which was in agreement with the data.

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 bending-torsional duct-induced swirl distortion on aerodynamic performance of a centrifugal compressor

NASA Astrophysics Data System (ADS)

Hou, Hongjuan; Wang, Leilei; Wang, Rui; Yang, Yanzhao

2017-04-01

A turbocharger compressor working in commercial vehicles, especially in some passenger cars, often works together with some pipes with complicated geometry as an air intake system, due to limit of available space in internal combustion engine compartments. These pipes may generate various distortions of physical parameters of the air at the inlet of the compressor and therefore the compressor aerodynamic performance deteriorates. Sometimes, the turbocharging engine fails to work at some operation points. This paper investigates the effects of various swirl distortions induced by different bending-torsional intake ducts on the aerodynamic performance of a turbocharger compressor by both 3D numerical simulations and experimental measurements. It was found that at the outlet of the pipes the different inlet ducts can generate different swirl distortions, twin vortices and bulk-like vortices with different rotating directions. Among them, the bulk-like vortices not only affect seriously the pressure distribution in the impeller domain, but also significantly deteriorate the compressor performance, especially at high flow rate region. And the rotating direction of the bulk-like vortices is also closely associated with the efficiency penalty. Besides the efficiency, the transient flow rate through a single impeller channel, or the asymmetric mass flow crossing the whole impeller, can be influenced by two disturbances. One is from the upstream bending-torsional ducts; other one is from the downstream volute.

Experimental evaluation of the effect of inlet distortion on compressor blade vibrations

NASA Technical Reports Server (NTRS)

Lubomski, J. F.

1979-01-01

Compressor rotor strain gage data from an engine test conducted with an inlet screen distortion were reduced and analyzed. These data are compared to data obtained from the same engine without inlet pressure distortion to determine the net effect of the distortion on the vibratory response of the compressor blades. The results obtained are presented.

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.

Flight evaluation of an engine static pressure noseprobe in an F-15 airplane

NASA Technical Reports Server (NTRS)

Foote, C. H.; Jaekel, R. F.

1981-01-01

The flight testing of an inlet static pressure probe and instrumented inlet case produced results consistent with sea-level and altitude stand testing. The F-15 flight test verified the basic relationship of total to static pressure ratio versus corrected airflow and automatic distortion downmatch with the engine pressure ratio control mode. Additionally, the backup control inlet case statics demonstrated sufficient accuracy for backup control fuel flow scheduling, and the station 6 manifolded production probe was in agreement with the flight test station 6 tota pressure probes.

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.

Summary of investigations of engine response to distorted inlet conditions

NASA Technical Reports Server (NTRS)

Biesiadny, T. J.; Braithwaite, W. M.; Soeder, R. H.; Abdelwahab, M.

1986-01-01

A survey is presented of experimental and analytical experience of the NASA Lewis Research Center in engine response to inlet temperature and pressure distortions. This includes a description of the hardware and techniques employed, and a summary of the highlights of experimental investigations and analytical modeling. Distortion devices successfully simulated inlet distortion, and knowledge was gained about compression system response to different types of distortion. A list of NASA research references is included.

Optimal Micro-Jet 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-jet secondary flow control is to demonstrate the viability and economy of Response Surface Methodology (RSM) to optimally design micro-jet 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 mass" micro-jet array designs. The term "low mass" micro-jet may refers to fluidic jets with total (integrated) mass flow ratios between 0.10 and 1.0 percent of the engine face mass flow. Therefore, this report examines optimal micro-jet array designs for compact inlets through a Response Surface Methodology.

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.

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.

Dynamic Distortion in a Short S-Shaped Subsonic Diffuser with Flow Separation. [Lewis 8 by 6 foot Supersonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Stumpf, R.; Neumann, H. E.; Giamati, C. C.

1983-01-01

An experimental investigation of the time varying distortion at the diffuser exit of a subscale HiMAT forebody and inlet was conducted at Mach 0.9 in the Lewis 8 by 6 foot Supersonic Wind Tunnel. A transitory separation was detected within the subsonic diffuser. Vortex generators were installed to eliminate the flow separation. Results from a study of the instantaneous pressure variations at the diffuser exit are presented. The time unsteady total pressures at the diffuser exit are computer interpolated and presented in the form of a movie showing the transitory separation. Limited data showing the instantaneous distortion levels is also presented.

Analysis of distortion data from TF30-P-3 mixed compression inlet test

NASA Technical Reports Server (NTRS)

King, R. W.; Schuerman, J. A.; Muller, R. G.

1976-01-01

A program was conducted to reduce and analyze inlet and engine data obtained during testing of a TF30-P-3 engine operating behind a mixed compression inlet. Previously developed distortion analysis techniques were applied to the data to assist in the development of a new distortion methodology. Instantaneous distortion techniques were refined as part of the distortion methodology development. A technique for estimating maximum levels of instantaneous distortion from steady state and average turbulence data was also developed as part of the program.

Numerical study of innovative scramjet inlets coupled to combustors using hydrocarbon-air mixture

NASA Astrophysics Data System (ADS)

Malo-Molina, Faure Joel

The research objective is to use high-fidelity multi-physics Computational Fluid Dynamics (CFD) analysis to characterize 3-D scramjet flowfields in two novel streamline traced circular configurations without axisymmetric profiles. This work builds on a body of research conducted over the past several years. In addition, this research provides the modeling and simulation support, prior to ground (wind tunnel) and flight experiment programs. Two innovative inlets, Jaws and Scoop, are analyzed and compared to a Baseline inlet, a current state of the art rectangular inlet used as a baseline for on/off-design conditions. The flight trajectory conditions selected were Mach 6 and a dynamic pressure of 1,500 psf (71.82 kPa), corresponding to a static pressure of 43.7 psf (2.09 kPa) and temperature of 400.8 R° (222.67 C°). All inlets are designed for equal flight conditions, equal contraction ratios and exit cross-sectional areas, thus facilitating their comparison and integration to a common combustor design. Analysis of these hypersonic inlets was performed to investigate distortion effects downstream in common generic combustors. These combustors include a single cavity acting as flame holder and strategically positioned fuel injection ports. This research not only seeks to identify the most successful integrated scramjet inlet/combustor design, but also investigates the flow physics and quantifies the integrated performance impact of the two novel scramjet inlet designs. It contributes to the hypersonic air-breathing community by providing analysis and predictions on directly-coupled combustor numerical experiments for developing pioneering inlets or nozzles for scramjets. Several validations and verifications of General Propulsion Analysis Chemical-kinetic and Two-phase (GPACT), the CFD tool, were conducted throughout the research. In addition, this study uses 13 gaseous species and 20 reactions for an Ethylene/air finite-rate chemical model. The key conclusions of this research are: (1) Flow distortion in the innovative inlets is similar to some of the distortion in the Baseline inlet, despite design differences. In both innovative inlets, the resulting flowfield distortions were due to shock boundary layer interactions similar to those found in the Baseline. The Baseline and Jaws performance attributes are stronger than Scoop, but Jaws accomplishes this while eradicating the cowl lip interaction, and lessening the total drag and spillage penalties. (2) The innovative inlets work best on-design, whereas for off-design, the traditional inlet yields a higher performance. Although the innovative inlets' designs mitigated some of the issues encountered in traditional configurations, they underperform at off-design conditions. The strategy used in Jaws was less prone to interaction with the near wall flow, and yields lesser pressure losses and higher efficiency at on-design conditions compared to the others. In general, the overall values for Scoop seem lowest of all due to lesser entrainment. Its drag coefficient and thrust to mass capture ratios are higher than the Baseline configuration. (3) Early pressure losses and flow distortions actually aid downstream combustion in all cases. Although interactions captured by the viscous simulations for the on-design conditions increase losses in the inlets, they enhance turbulence in the isolator, favoring the mixing of air and fuel, and improving the overall factor of the system. Jaws inlet demonstrates the most valuable design with higher performance, but its factor later in the combustor drops relative to its rectangular counterpart. (4) A parametric study of the location and direction of injection is conducted to select the configuration for fuel penetration, mixing factor (factor) and other combustion qualities. Although the trends observed with and without chemical reactions are the same, the former yields roughly 10% higher mixing factor. Unlike at frozen conditions, when chemical reactions are considered, a high compression area was observed upstream of the cavity, not present when modeling Jaws. The upstream reactions from the cavity have a significant impact on the development of the shear layers and downstream development of the entire combustion. (5) Steady and unsteady simulations are conducted to characterize the ignition process, flame anchoring and flashback effects. This unsteadiness enlarges the circulation region in and around the cavity, allowing the reactions to propagate forward through the shear layer, and increases the mixing factor. In Scoop, these effects are exacerbated due to the thicker low energy profile surrounding the walls and most of the lower section of the combustor. In the steady assumptions, the forward reactions and their effects are positioned farthest upstream, closest to the combustor entrance. (6) Unsteady Reynolds Average Navier-Stokes (URANS) and Large Eddy Simulation (LES) modeling are compared to explore overall flow structure and for comparison of individual numerical methods. In URANS, the flashback effects are midway between the entrance and the step, whereas in LES, this effect is near the edge of the step in addition to yielding a higher combustion factor. Thus, the turbulence model and inflow assumptions can critically affect the total outcome of such devices.

F100(3) parallel compressor computer code and user's manual

NASA Technical Reports Server (NTRS)

Mazzawy, R. S.; Fulkerson, D. A.; Haddad, D. E.; Clark, T. A.

1978-01-01

The Pratt & Whitney Aircraft multiple segment parallel compressor model has been modified to include the influence of variable compressor vane geometry on the sensitivity to circumferential flow distortion. Further, performance characteristics of the F100 (3) compression system have been incorporated into the model on a blade row basis. In this modified form, the distortion's circumferential location is referenced relative to the variable vane controlling sensors of the F100 (3) engine so that the proper solution can be obtained regardless of distortion orientation. This feature is particularly important for the analysis of inlet temperature distortion. Compatibility with fixed geometry compressor applications has been maintained in the model.

Application of computational fluid dynamics to the study of vortex flow control for the management of inlet distortion

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Gibb, James

1992-01-01

A study is presented to demonstrate that the Reduced Navier-Stokes code RNS3D can be employed effectively to develop a vortex generator installation that minimizes engine face circumferential distortion by controlling the development of secondary flow. The necessary computing times are small enough to show that similar studies are feasible within an analysis-design environment with all its constraints of costs and time. This study establishes the nature of the performance enhancements that can be realized with vortex flow control, and indicates a set of aerodynamic properties that can be utilized to arrive at a successful vortex generator installation design.

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.

Response of a small-turboshaft-engine compression system to inlet temperature distortion

NASA Technical Reports Server (NTRS)

Biesiadny, T. J.; Klann, G. A.; Little, J. K.

1984-01-01

An experimental investigation was conducted into the response of a small-turboshaft-engine compression system to steady-state and transient inlet temperature distortions. Transient temperature ramps range from less than 100 K/sec to above 610 K/sec and generated instantaneous temperatures to 420 K above ambient. Steady-state temperature distortion levels were limited by the engine hardware temperature list. Simple analysis of the steady-state distortion data indicated that a particle separator at the engine inlet permitted higher levels of temperature distortion before onset of compressor surge than would be expected without the separator.

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

NASA Technical Reports Server (NTRS)

Kim, Hyun D.; Frate, Franco C.

2001-01-01

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

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.

Mitigation of Engine Inlet Distortion Through Adjoint-Based Design

NASA Technical Reports Server (NTRS)

Ordaz, Irian; Rallabhandi, Sriram; Nielsen, Eric J.; Diskin, Boris

2017-01-01

The adjoint-based design capability in FUN3D is extended to allow efficient gradient- based optimization and design of concepts with highly integrated aero-propulsive systems. A circumferential distortion calculation, along with the derivatives needed to perform adjoint-based design, have been implemented in FUN3D. This newly implemented distortion calculation can be used not only for design but also to drive the existing mesh adaptation process and reduce the error associated with the fan distortion calculation. The design capability is demonstrated by the shape optimization of an in-house aircraft concept equipped with an aft fuselage propulsor. The optimization objective is the minimization of flow distortion at the aerodynamic interface plane of this aft fuselage propulsor.

Effect of inlet ingestion of a wing tip vortex on compressor face flow and turbojet stall margin

NASA Technical Reports Server (NTRS)

Mitchell, G. A.

1975-01-01

A two-dimensional inlet was alternately mated to a coldpipe plug assembly and a J85-GE-13 turbojet engine, and placed in a Mach 0.4 stream so as to ingest the tip vortex of a forward mounted wing. Vortex properties were measured just forward of the inlet and at the compressor face. Results show that ingestion of a wing tip vortex by a turbojet engine can cause a large reduction in engine stall margin. The loss in stall compressor pressure ratio was primarily dependent on vortex location and rotational direction and not on total-pressure distortion.

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.)

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

NASA Technical Reports Server (NTRS)

Slater, John W.

2014-01-01

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

Impact of inlet coherent motions on compressor performance

NASA Astrophysics Data System (ADS)

Forlese, Jacopo; Spoleti, Giovanni

2017-08-01

Automotive engine induction systems may be characterized by significant flow angularity and total pressure distortion at the compressor inlet. The impact of the swirl on compressor performance should be quantified to guide the design of the induction systems. In diesel engines, the presence of a valve for flow reduction and control of low pressure EGR recirculation could generate coherent motion and influence the performance of the compressor. Starting from experimental map, the compressor speed-lines have been simulated using a 3D CFD commercial code imposing different concept motion at the inlet. The swirl intensity, the direction and the number of vortices have been imposed in order to taking into account some combinations. Finally, a merit function has been defined to evaluate the performance of the compressor with the defined swirl concepts. The aim of the current work is to obtain an indication on the effect of a swirling motion at the compressor inlet on the engine performance and provide a guideline to the induction system design.

Comparison of theoretical and experimental boundary-layer development in a Mach 2.5 mixed-compression inlet

NASA Technical Reports Server (NTRS)

Hingst, W. R.; Towne, C. E.

1974-01-01

An analytical investigation was made of the boundary layer flow in an axisymmetric Mach 2.5 mixed compression inlet, and the results were compared with experimental measurements. The inlet tests were conducted in the Lewis 10- by 10-foot supersonic wind tunnel at a unit Reynolds number of 8.2 million/m. The inlet incorporated porous bleed regions for boundary layer control, and the effect of this bleed was taken into account in the analysis. The experimental boundary layer data were analyzed by using similarity laws from which the skin friction coefficient was obtained. The boundary layer analysis included predictions of laminar and turbulent boundary layer growth, transition, and the effects of the shock boundary layer interactions. In addition, the surface static pressures were compared with those obtained from an inviscid characteristics program. The results of investigation showed that the analytical techniques gave satisfactory predictions of the boundary layer flow except in regions that were badly distorted by the terminal shock.

Grid Resolution Study over Operability Space for a Mach 1.7 Low Boom External Compression Inlet

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.

2014-01-01

This paper presents a statistical methodology whereby the probability limits associated with CFD grid resolution of inlet flow analysis can be determined which provide quantitative information on the distribution of that error over the specified operability range. The objectives of this investigation is to quantify the effects of both random (accuracy) and systemic (biasing) errors associated with grid resolution in the analysis of the Lockheed Martin Company (LMCO) N+2 Low Boom external compression supersonic inlet. The study covers the entire operability space as defined previously by the High Speed Civil Transport (HSCT) High Speed Research (HSR) program goals. The probability limits in terms of a 95.0% confidence interval on the analysis data were evaluated for four ARP1420 inlet metrics, namely (1) total pressure recovery (PFAIP), (2) radial hub distortion (DPH/P), (3) ) radial tip distortion (DPT/P), and (4) ) circumferential distortion (DPC/P). In general, the resulting +/-0.95 delta Y interval was unacceptably large in comparison to the stated goals of the HSCT program. Therefore, the conclusion was reached that the "standard grid" size was insufficient for this type of analysis. However, in examining the statistical data, it was determined that the CFD analysis results at the outer fringes of the operability space were the determining factor in the measure of statistical uncertainty. Adequate grids are grids that are free of biasing (systemic) errors and exhibit low random (precision) errors in comparison to their operability goals. In order to be 100% certain that the operability goals have indeed been achieved for each of the inlet metrics, the Y+/-0.95 delta Y limit must fall inside the stated operability goals. For example, if the operability goal for DPC/P circumferential distortion is =0.06, then the forecast Y for DPC/P plus the 95% confidence interval on DPC/P, i.e. +/-0.95 delta Y, must all be less than or equal to 0.06.

Validation of the Predicted Circumferential and Radial Mode Sound Power Levels in the Inlet and Exhaust Ducts of a Fan Ingesting Distorted Inflow

NASA Technical Reports Server (NTRS)

Koch, L. Danielle

2012-01-01

Fan inflow distortion tone noise has been studied computationally and experimentally. Data from two experiments in the NASA Glenn Advanced Noise Control Fan rig have been used to validate acoustic predictions. The inflow to the fan was distorted by cylindrical rods inserted radially into the inlet duct one rotor chord length upstream of the fan. The rods were arranged in both symmetric and asymmetric circumferential patterns. In-duct and farfield sound pressure level measurements were recorded. It was discovered that for positive circumferential modes, measured circumferential mode sound power levels in the exhaust duct were greater than those in the inlet duct and for negative circumferential modes, measured total circumferential mode sound power levels in the exhaust were less than those in the inlet. Predicted trends in overall sound power level were proven to be useful in identifying circumferentially asymmetric distortion patterns that reduce overall inlet distortion tone noise, as compared to symmetric arrangements of rods. Detailed comparisons between the measured and predicted radial mode sound power in the inlet and exhaust duct indicate limitations of the theory.

High-tip-speed, low-loading transonic fan stage. Part 3: Final report

NASA Technical Reports Server (NTRS)

Ware, T. C.; Kobayashi, R. J.; Jackson, R. J.

1974-01-01

Tests were conducted on a high-tip-speed, low-loading transonic fan stage to determine the performance and inlet flow distortion tolerance of the design. The fan was designed for high efficiency at a moderate pressure ratio by designing the hub section to operate at minimum loss when the tip operates with an oblique shock. The design objective was an efficiency of 86 percent at a pressure ratio of 1.5, a specific flow (flow per unit annulus area) of 42 lb/sec-sq. ft (205.1 kgm/sec-m sq), and a tip speed of 1600 ft/sec (488.6 m/sec). During testing, a peak efficiency of 84 percent was achieved at design speed and design specific flow. At the design speed and pressure ratio, the flow was 4 percent greater than design, efficiency was 81 percent, and a stall margin of 24 percent was obtained. The stall line was improved with hub radial distortion but was reduced when the stage was tested with tip radial and circumferential flow distortions. Blade-to-blade values of static pressures were measured over the rotor blade tips.

Forced Response Analysis of a Fan with Boundary Layer Inlet Distortion

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Reddy, T. S. R.; Coroneos, Rula M.

2014-01-01

Boundary layer ingesting propulsion systems have the potential to significantly reduce fuel burn for future generations of commercial aircraft, but these systems must be designed to overcome the challenge of high dynamic stresses in fan blades due to forced response. High dynamic stresses can lead to high cycle fatigue failures. High-fidelity computational analysis of the fan aeromechanics is integral to an ongoing effort to design a boundary layer ingesting inlet and fan for a wind-tunnel test. An unsteady flow solution from a Reynoldsaveraged Navier Stokes analysis of a coupled inlet-fan system is used to calculate blade unsteady loading and assess forced response of the fan to distorted inflow. Conducted prior to the mechanical design of a fan, the initial forced response analyses performed in this study provide an early look at the levels of dynamic stresses that are likely to be encountered. For the boundary layer ingesting inlet, the distortion contains strong engine order excitations that act simultaneously. The combined effect of these harmonics was considered in the calculation of the forced response stresses. Together, static and dynamic stresses can provide the information necessary to evaluate whether the blades are likely to fail due to high cycle fatigue. Based on the analyses done, the overspeed condition is likely to result in the smallest stress margin in terms of the mean and alternating stresses. Additional work is ongoing to expand the analyses to off-design conditions, on-resonance conditions, and to include more detailed modeling of the blade structure.

F-15 inlet/engine test techniques and distortion methodologies studies. Volume 2: Time variant data quality analysis plots

NASA Technical Reports Server (NTRS)

Stevens, C. H.; Spong, E. D.; Hammock, M. S.

1978-01-01

Time variant data quality analysis plots were used to determine if peak distortion data taken from a subscale inlet model can be used to predict peak distortion levels for a full scale flight test vehicle.

Laser Displacement Measurements of Fan Blades in Resonance and Flutter During the Boundary Layer Ingesting Inlet and Distortion-Tolerant Fan Test

NASA Technical Reports Server (NTRS)

Duffy, Kirsten P.; Provenza, Andrew J.; Bakhle, Milind A.; Min, James B.; Abdul-Aziz, Ali

2018-01-01

NASA's Advanced Air Transport Technology Project is investigating boundary layer ingesting propulsors for future subsonic commercial aircraft to improve aircraft efficiency, thereby reducing fuel burn. To that end, a boundary layer ingesting inlet and distortion-tolerant fan stage was designed, fabricated, and tested within the 8' x 6' Supersonic Wind Tunnel at NASA Glenn Research Center. Because of the distortion in the air flow over the fan, the blades were designed to withstand a much higher aerodynamic forcing than for a typical clean flow. The blade response for several resonance modes were measured during start-up and shutdown, as well as at near 85% design speed. Flutter in the first bending mode was also observed in the fan at the design speed, at an off-design condition, although instabilities were difficult to instigate with this fan in general. Blade vibrations were monitored through twelve laser displacement probes that were placed around the inner circumference of the casing, at the blade leading and trailing edges. These probes captured the movement of all the blades during the entire test. Results are presented for various resonance mode amplitudes, frequencies and damping, as well as flutter amplitudes and frequency. Benefits and disadvantages of laser displacement probe measurements versus strain gage measurements are discussed.

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.

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.

Theoretical studies of tone noise from a fan rotor

NASA Technical Reports Server (NTRS)

Rao, G. V. R.; Chu, W. T.; Digumarthi, R. V.

1973-01-01

An analytical study was made of some possible rotor alone noise sources of dipole, quadrapole and monopole characters which generate discrete tone noise. Particular emphasis is given to the tone noise caused by fan inlet flow distortion and turbulence. Analytical models are developed to allow prediction of absolute levels. Experimental data measured on a small scale fan is presented which indicates inlet turbulence interaction with a fan rotor can be a source of tone noise. Predicted and measured tone noise for the small scale rotor are shown to be in reasonable agreement.

Data analysis of P sub T/P sub S noseboom probe testing on F100 engine P680072 at NASA Lewis Research Center

NASA Technical Reports Server (NTRS)

Foote, C. H.

1980-01-01

Results from the altitude testing of a P sub T/P sub S noseboom probe on the F100 engine are discused. The results are consistent with sea level test results. The F100 engine altitude test verified automatic downmatch with the engine pressure ratio control, and backup control inlet case static pressure demonstrated sufficient accuracy for backup control fuel flow scheduling. The production P6 probe measured Station 6 pressures accurately for both undistorted and distorted inlet airflows.

Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Aircraft

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.; Conners, Timothy R.; Wayman, Thomas R.

2009-01-01

A computational analysis of a Gulfstream isentropic external compression supersonic inlet coupled to a Rolls-Royce fan was completed. The inlet was designed for a small, low sonic boom supersonic vehicle with a design cruise condition of M = 1.6 at 45,000 feet. The inlet design included an annular bypass duct that routed flow subsonically around an engine-mounted gearbox and diverted flow with high shock losses away from the fan tip. Two Reynolds-averaged Navier-Stokes codes were used for the analysis: an axisymmetric code called AVCS for the inlet and a 3-D code called SWIFT for the fan. The codes were coupled at a mixing plane boundary using a separate code for data exchange. The codes were used to determine the performance of the inlet/fan system at the design point and to predict the performance and operability of the system over the flight profile. At the design point the core inlet had a recovery of 96 percent, and the fan operated near its peak efficiency and pressure ratio. A large hub radial distortion generated in the inlet was not eliminated by the fan and could pose a challenge for subsequent booster stages. The system operated stably at all points along the flight profile. Reduced stall margin was seen at low altitude and Mach number where flow separated on the interior lips of the cowl and bypass ducts. The coupled analysis gave consistent solutions at all points on the flight profile that would be difficult or impossible to predict by analysis of isolated components.

Coupled Analysis of an Inlet and Fan for a Quiet Supersonic Jet

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.; Conners, Timothy R.; Wayman, Thomas R.

2010-01-01

A computational analysis of a Gulfstream isentropic external compression supersonic inlet coupled to a Rolls-Royce fan has been completed. The inlet was designed for a small, low sonic boom supersonic vehicle with a design cruise condition of M = 1.6 at 45,000 ft. The inlet design included an annular bypass duct that routed flow subsonically around an engine-mounted gearbox and diverted flow with high shock losses away from the fan tip. Two Reynolds-averaged Navier-Stokes codes were used for the analysis: an axisymmetric code called AVCS for the inlet and a three dimensional (3-D) code called SWIFT for the fan. The codes were coupled at a mixing plane boundary using a separate code for data exchange. The codes were used to determine the performance of the inlet/fan system at the design point and to predict the performance and operability of the system over the flight profile. At the design point the core inlet had a recovery of 96 percent, and the fan operated near its peak efficiency and pressure ratio. A large hub radial distortion generated in the inlet was not eliminated by the fan and could pose a challenge for subsequent booster stages. The system operated stably at all points along the flight profile. Reduced stall margin was seen at low altitude and Mach number where flow separated on the interior lips of the cowl and bypass ducts. The coupled analysis gave consistent solutions at all points on the flight profile that would be difficult or impossible to predict by analysis of isolated components.

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.

A comparative study of Full Navier-Stokes and Reduced Navier-Stokes analyses for separating flows within a diffusing inlet S-duct

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Reddy, D. R.; Kapoor, K.

1993-01-01

A three-dimensional implicit Full Navier-Stokes (FNS) analysis and a 3D Reduced Navier Stokes (RNS) initial value space marching solution technique has been applied to a class of separated flow problems within a diffusing S-duct configuration characterized by vortex-liftoff. Both the FNS and the RNS solution technique were able to capture the overall flow physics of vortex lift-off, and gave remarkably similar results which agreed reasonably well with the experimental measured averaged performance parameters of engine face total pressure recovery and distortion. However, the Full Navier-Stokes and Reduced Navier-Stokes also consistently predicted separation further downstream in the M2129 inlet S-duct than was indicated by experimental data, thus compensating errors were present in the two Navier-Stokes analyses. The difficulties encountered in the Navier-Stokes separations analyses of the M2129 inlet S-duct center primarily on turbulence model issues, and these focused on two distinct but different phenomena, namely, (1) characterization of low skin friction adverse pressure gradient flows, and (2) description of the near wall behavior of flows characterized by vortex lift-off.

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

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

DTIC Science & Technology

2013-01-01

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

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.

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.

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.

Low-Cost, User-Friendly, Rapid Analysis of Dynamic Data System Established

NASA Technical Reports Server (NTRS)

Arend, David J.

2004-01-01

An issue of primary importance to the development of new jet and certain other airbreathing combined-cycle powered aircraft is the advancement of airframe-integrated propulsion technologies. Namely, engine inlets and their systems and subsystems are required to capture, convert, and deliver the atmospheric airflow demanded by such engines across their operating envelope in a form that can be used to provide efficient, stable thrust. This must be done while also minimizing aircraft drag and weight. Revolutionary inlet designs aided by new technologies are needed to enable new missions. An unwanted byproduct of pursuing these inlet technologies is increased time-variant airflow distortion. Such distortions reduce propulsion system stability, performance, operability, and life. To countermand these limitations and fully evaluate the resulting configurations, best practices dictate that this distortion be experimentally measured at large scale and analyzed. The required measurements consist of those made by an array of high-response pressure transducers located in the flow field at the aerodynamic interface plane (AIP) between the inlet and engine. Although the acquisition of the necessary pitot-pressure time histories is relatively straight-forward, until recent years, the analysis has proved to be very time-consuming, tedious, and expensive. To transform the analysis of these data into a tractable and timely proposition, researchers at the NASA Glenn Research Center created and established the Rapid Analysis of Dynamic Data (RADD) system. The system provides complete, near real-time analysis of time-varying inlet airflow distortion datasets with report quality output. This fully digital approach employs Institute of Electrical and Electronics Engineers (IEEE) binary data file format standardization to establish data-acquisition-system-independent processing on low cost personal computers. Features include invalid instrumentation code-out, logging, and multiple replacement schemes as needed for each channel of instrumentation. The AIP pressure distribution can be interpolated to simulate measurements by alternate AIP probe arrays, if desired. In addition, the RADD system provides for the application of filters that can be used to focus the analysis on the frequency range of interest.

The effect of unsteadiness on the time-mean thermal loads in a turbine stage

NASA Technical Reports Server (NTRS)

Kirtley, K. R.; Celestina, M. L.; Adamczyk, J. J.

1993-01-01

Two steady numerical analysis methods and one unsteady method are used to study the viscous three-dimensional flow in the middle stage of the Pratt & Whitney alternate design Space Shuttle Main Engine fuel turbine. The principal characteristic of this flow is that the secondary flows generated in the rotor blade reconfigure a radial inlet total temperature distortion into one with a pitchwise exit hot streak distortion. Secondary flows in the following vane redistribute the radial variation while unsteadiness causes a segregation of hot and cold flow from the hot streak within the vane. Such redistribution and segregation can lead to unexpected thermal loads and reduced durability. The physical phenomena and the ability of a steady analysis to capture them are investigated by performing a numerical experiment whereby the results of the two steady analysis methods are compared to the time-mean of the unsteady simulation. The flow physics related to the segregation and mixing of total temperature are discussed.

Hybrid Manipulation of Streamwise Vorticity in a Diffuser Boundary Layer

NASA Astrophysics Data System (ADS)

Gissen, Abraham; Vukasinovic, Bojan; Culp, John; Glezer, Ari

2010-11-01

The formation of streamwise vorticity concentrations by exploiting the interaction of surface-mounted passive (micro-vanes) and active (synthetic jets) flow control elements with the cross flow is investigated experimentally in a small-scale serpentine duct at high subsonic speeds (up to M = 0.6). Streamwise vortices can be a key element in the mitigation of the adverse effects on pressure recovery and distortion caused by the naturally occurring secondary flows in embedded propulsion systems with complex inlet geometries. Counter rotating and single-sense vortices are formed using conventional passive micro-vanes and active high-power synthetic jet actuators. Interaction of the flow control elements is examined through a hybrid actuation scheme whereby synthetic jet actuation augments the primary vanes' vortices resulting in dynamic enhancement of their strength. It is shown that such sub-boundary layer individual vortices can merge and evolve into duct-scale vortical structures that counteract the inherent secondary flow and mitigates global flow distortion.

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

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.

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

NASA Technical Reports Server (NTRS)

Frederick, Michael; Ratnayake, Nalin

2011-01-01

The results are described of the Rake Airflow Gage Experiment (RAGE), which was designed and fabricated to support the flight test of a new supersonic inlet design using Dryden's Propulsion Flight Test Fixture (PFTF) and F-15B testbed airplane (see figure). The PFTF is a unique pylon that was developed for flight-testing propulsion-related experiments such as inlets, nozzles, and combustors over a range of subsonic and supersonic flight conditions. The objective of the RAGE program was to quantify the local flowfield at the aerodynamic interface plane of the Channeled Centerbody Inlet Experiment (CCIE). The CCIE is a fixed representation of a conceptual mixed-compression supersonic inlet with a translating biconic centerbody. The primary goal of RAGE was to identify the relationship between free-stream and local Mach number in the low supersonic regime, with emphasis on the identification of the particular free-stream Mach number that produced a local Mach number of 1.5. Measurements of the local flow angularity, total pressure distortion, and dynamic pressure over the interface plane were also desired. The experimental data for the RAGE program were obtained during two separate research flights. During both flights, local flowfield data were obtained during straight and level acceleration segments out to steady-state test points. The data obtained from the two flights showed small variations in Mach number, flow angularity, and dynamic pressure across the interface plane at all flight conditions. The data show that a free-stream Mach number of 1.65 will produce the desired local Mach number of 1.5 for CCIE. The local total pressure distortion over the interface plane at this condition was approximately 1.5%. At this condition, there was an average of nearly 2 of downwash over the interface plane. This small amount of downwash is not expected to adversely affect the performance of the CCIE inlet.

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

NASA Astrophysics Data System (ADS)

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

2015-08-01

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

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.

Active control of Boundary Layer Separation & Flow Distortion in Adverse Pressure Gradient Flows via Supersonic Microjets

NASA Technical Reports Server (NTRS)

Alvi, Farrukh S.; Gorton, Susan (Technical Monitor)

2005-01-01

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 aim of this research has been to understand the nature of separation and more importantly, to explore techniques to actively control this flow separation. In particular, the use of supersonic microjets as a means of controlling boundary layer separation was explored. The geometry used for the early part of this study was a simple diverging Stratford ramp, equipped with arrays of supersonic microjets. Initial results, based on the mean surface pressure distribution, surface flow visualization and Planar Laser Scattering (PLS) indicated a reverse flow region. We implemented supersonic microjets to control this separation and flow visualization results appeared to suggest that microjets have a favorable effect, at least to a certain extent. However, the details of the separated flow field were difficult to determine based on surface pressure distribution, surface flow patterns and PLS alone. It was also difficult to clearly determine the exact influence of the supersonic microjets on this flow. In the latter part of this study, the properties of this flow-field and the effect of supersonic microjets on its behavior were investigated in further detail using 2-component (planar) Particle Image Velocimetry (PIV). The results clearly show that the activation of microjets eliminated flow separation and resulted in a significant increase in the momentum of the fluid near the ramp surface. Also notable is the fact that the gain in momentum due to the elimination of flow separation is at least an order of magnitude larger (two orders of magnitude larger in most cases) than the momentum injected by the microjets and is accomplished with very little mass flow through the microjets.

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.

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.

Task 7: Endwall treatment inlet flow distortion analysis

NASA Technical Reports Server (NTRS)

Hall, E. J.; Topp, D. A.; Heidegger, N. J.; McNulty, G. S.; Weber, K. F.; Delaney, R. A.

1996-01-01

The overall objective of this study was to develop a 3-D numerical analysis for compressor casing treatment flowfields, and to perform a series of detailed numerical predictions to assess the effectiveness of various endwall treatments for enhancing the efficiency and stall margin of modern high speed fan rotors. Particular attention was given to examining the effectiveness of endwall treatments to counter the undesirable effects of inflow distortion. Calculations were performed using three different gridding techniques based on the type of casing treatment being tested and the level of complexity desired in the analysis. In each case, the casing treatment itself is modeled as a discrete object in the overall analysis, and the flow through the casing treatment is determined as part of the solution. A series of calculations were performed for both treated and untreated modern fan rotors both with and without inflow distortion. The effectiveness of the various treatments were quantified, and several physical mechanisms by which the effectiveness of endwall treatments is achieved are discussed.

Flow control in a diffusing S-Duct

NASA Technical Reports Server (NTRS)

Vakili, A. D.; Wu, J. M.; Liver, P.; Bhat, M. K.

1985-01-01

Accurate measurements have been made of secondary flow in a 1.51 area ratio diffusing 30 deg - 30 deg S-Duct with circulair cross section. Turbulent flow was entering the duct at Mach number of 0.6, the boundary layer thickness at the duct entrance was ten percent of the duct inlet diameter. Through measurements made, local flow velocity vector as well as static and total pressures mapping of the flow at several stations were obtained. Strong secondary flow was measured in the first bend which continued into the second bend with new vorticity produced in there in the opposite direction. Surface oil flow visualization and wall pressures indicated a region of separated flow starting at theta approximately equal to 22 deg on the inside of the first bend up to theta approximately equal to 44 deg on the outside of the second bend. The flow separated in 'cyclone' form and never reattached in the duct. As a result of the secondary flow and the flow separation, significant total pressure distortion was observed at the exit of the duct. Using flow control devices the separation was eliminated while the exit distortion was improved.

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.

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.

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.

Effect of number of probes and their orientation on the calculation of several compressor face distortion descriptors

NASA Technical Reports Server (NTRS)

Stoll, F.; Tremback, J. W.; Arnaiz, H. H.

1979-01-01

A study was performed to determine the effects of the number and position of total pressure probes on the calculation of five compressor face distortion descriptors. This study used three sets of 320 steady state total pressure measurements that were obtained with a special rotating rake apparatus in wind tunnel tests of a mixed-compression inlet. The inlet was a one third scale model of the inlet on a YF-12 airplane, and it was tested in the wind tunnel at representative flight conditions at Mach numbers above 2.0. The study shows that large errors resulted in the calculation of the distortion descriptors even with a number of probes that were considered adequate in the past. There were errors as large as 30 and -50 percent in several distortion descriptors for a configuration consisting of eight rakes with five equal-area-weighted probes on each rake.

F-15 inlet/engine test techniques and distortion methodologies studies. Volume 1: Technical discussion

NASA Technical Reports Server (NTRS)

Stevens, C. H.; Spong, E. D.; Hammock, M. S.

1978-01-01

Peak distortion data taken from a subscale inlet model were studied to determine if the data can be used to predict peak distortion levels for a full scale flight test vehicle, and to provide a better understanding of the time variant total pressure distortion and the attendant effects of Reynolds number/scale and frequency content. The data base used to accomplish this goal covered a range from Mach 0.4 to 2.5 and an angle of attack range from -10 degrees to +12 degrees. Data are presented which show that: (1) increasing the Reynolds number increases total pressure recovery, decreases peak distortion, and decreases turbulence, (2) increasing the filter cutoff frequency increases both peak distortion and turbulence, and (3) the effect of engine presence on total pressure recovery, peak distortion, and turbulence is small but favorable.

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.

Comparison of steady and unsteady secondary flows in a turbine stator cascade

NASA Technical Reports Server (NTRS)

Hebert, Gregory J.; Tiederman, William G.

1989-01-01

The effect of periodic rotor wakes on the secondary flow structure in a turbine stator cascade was investigated. A mechanism simulated the wakes shed from rotor blades by passing cylindrical rods across the inlet to a linear cascade installed in a recirculating water flow loop. Velocity measurements showed a passage vortex, similar to that seen in steady flow, during the time associated with undisturbed fluid. However, as the rotor wake passed through the blade row, a large crossflow toward the suction surface was observed in the midspan region. This caused the development of two large areas of circulation between the midspan and endwall regions, significantly distorting and weakening the passage vortices.

Linearized blade row compression component model. Stability and frequency response analysis of a J85-3 compressor

NASA Technical Reports Server (NTRS)

Tesch, W. A.; Moszee, R. H.; Steenken, W. G.

1976-01-01

NASA developed stability and frequency response analysis techniques were applied to a dynamic blade row compression component stability model to provide a more economic approach to surge line and frequency response determination than that provided by time-dependent methods. This blade row model was linearized and the Jacobian matrix was formed. The clean-inlet-flow stability characteristics of the compressors of two J85-13 engines were predicted by applying the alternate Routh-Hurwitz stability criterion to the Jacobian matrix. The predicted surge line agreed with the clean-inlet-flow surge line predicted by the time-dependent method to a high degree except for one engine at 94% corrected speed. No satisfactory explanation of this discrepancy was found. The frequency response of the linearized system was determined by evaluating its Laplace transfer function. The results of the linearized-frequency-response analysis agree with the time-dependent results when the time-dependent inlet total-pressure and exit-flow function amplitude boundary conditions are less than 1 percent and 3 percent, respectively. The stability analysis technique was extended to a two-sector parallel compressor model with and without interstage crossflow and predictions were carried out for total-pressure distortion extents of 180 deg, 90 deg, 60 deg, and 30 deg.

Active Control of Jet Engine Inlet Flows

DTIC Science & Technology

2007-03-31

These S-shaped ducts do not provide a direct line of sight to the compressor blades , thus hiding the engine from incoming radar waves. Also, serpentine...circumferential distortion pattern acts as an unsteady forcing function, inducing blade vibration that can result in structural fatigue and failure 3. This...shortcoming occurs when the rotor blades pass through regions of reduced axial velocity (i.e., where the total pressure is low). In these areas, since the

Temperature distortion generator for turboshaft engine testing

NASA Technical Reports Server (NTRS)

Klann, G. A.; Barth, R. L.; Biesiadny, T. J.

1984-01-01

The procedures and unique hardware used to conduct an experimental investigation into the response of a small-turboshaft-engine compression system to various hot gas ingestion patterns are presented. The temperature distortion generator described herein uses gaseous hydrogen to create both steady-state and time-variant, or transient, temperature distortion at the engine inlet. The range of transient temperature ramps produced by the distortion generator during the engine tests was from less than 111 deg K/sec (200 deg R/sec) to above 611 deg K/sec (1100 deg R/sec); instantaneous temperatures to 422 deg K (760 deg R) above ambient were generated. The distortion generator was used to document the maximum inlet temperatures and temperature rise rates that the compression system could tolerate before the onset of stall for various circumferential distortions as well as the compressor system response during stall.

Inlet Trade Study for a Low-Boom Aircraft Demonstrator

NASA Technical Reports Server (NTRS)

Heath, Christopher M.; Slater, John W.; Rallabhandi, Sriram K.

2016-01-01

Propulsion integration for low-boom supersonic aircraft requires careful inlet selection, placement, and tailoring to achieve acceptable propulsive and aerodynamic performance, without compromising vehicle sonic boom loudness levels. In this investigation, an inward-turning streamline-traced and axisymmetric spike inlet are designed and independently installed on a conceptual low-boom supersonic demonstrator aircraft. The airframe was pre-shaped to achieve a target ground under-track loudness of 76.4 PLdB at cruise using an adjoint-based design optimization process. Aircraft and inlet performance characteristics were obtained by solution of the steady-state Reynolds-averaged Navier-Stokes equations. Isolated cruise inlet performance including total pressure recovery and distortion were computed and compared against installed inlet performance metrics. Evaluation of vehicle near-field pressure signatures, along with under- and off-track propagated loudness levels is also reported. Results indicate the integrated axisymmetric spike design offers higher inlet pressure recovery, lower fan distortion, and reduced sonic boom. The vehicle with streamline-traced inlet exhibits lower external wave drag, which translates to a higher lift-to-drag ratio and increased range capability.

Internal flows and force matrices in axial flow inducers

NASA Astrophysics Data System (ADS)

Bhattacharyya, Abhijit

1994-01-01

Axial flow inducers such as those used in high speed rocket engine turbopumps are subject to complex internal flows and fluid-induced lateral and rotordynamic forces. An investigation of these internal flows was conducted using boundary layer flow visualization on the blades, hub and housing of unshrouded and shrouded inducers. Results showed that the blade boundary layer flows have strong radial components at off-design conditions and remain attached to the blade surface at all flow coefficients tested. The origin of upstream swirling backflow was found to be at the discharge plane of the inducer. In addition, flow reversal was observed at the suction side blade tip near the leading edge in a shrouded inducer. Re-entry of the hub boundary layer flow, a downstream backflow, into the blade passage area was observed at flow coefficients below design. For unshrouded inducers the radially outward flow near the blade tip mixed with the leakage flow to form the upstream backflow. The lateral and rotordynamic forces acting on an inducer due to an imposed whirl motion was also investigated at various flow coefficients. It was found that the rotordynamic force data at various whirl frequency ratios does not allow a normal quadratic fit; consequently the conventional inertial, stiffness and damping coefficients cannot be obtained and a definite whirl ratio describing the instability region does not result. Application of an actuator disk theory proved to be inaccurate in estimating the rotordynamic tangential force in a non-whirling inducer. The effect of upstream and downstream flow distortions on the rotordynamic and lateral forces on an inducer were studied. It was found that at flow coefficients below design, large lateral forces occurred in the presence of a downstream asymmetry. Results of inlet distortion experiments show that a strong inlet shear causes a significant increase in the lateral force. Cavitation was found to have important consequences for fluid-induced rotordynamic forces. These forces become destabilizing for both forward and reverse whirl. Decreasing cavitation numbers caused an increase in the magnitudes of the destabilizing forces.

Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- x 15-foot low speed wind tunnel

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Neiner, George; Bencic, Timothy J.; Flood, Joseph D.; Amuedo, Kurt C.; Strock, Thomas W.

1990-01-01

A 9.2 percent scale Short Takeoff and Vertical Landing (STOVL) hot gas ingestion model was designed and built by McDonnell Douglas Corporation (MCAIR) and tested in the Lewis Research Center 9 x 15 foot Low Speed Wind Tunnel (LSWT). Hot gas ingestion, the entrainment of heated engine exhaust into the inlet flow field, is a key development issure for advanced short takeoff and vertical landing aircraft. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The model support system had four degrees of freedom - pitch, roll, yaw, and vertical height variation. The model support system also provided heated high-pressure air for nozzle flow and a suction system exhaust for inlet flow. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Test and data analysis results from Phase 2 and flow visualization from both Phase 1 and 2 are documented. A description of the model and facility modifications is also provided. Headwind velocity was varied from 10 to 23 kn. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R. These results will contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours, the model airframe heating, and the location of the ground flow separation.

Experimental quiet engine program aerodynamic performance of fan A

NASA Technical Reports Server (NTRS)

Giffin, R. G.; Parker, D. E.; Dunbar, L. W.

1971-01-01

The aerodynamic component test results are presented of fan A, one of two high-bypass-ratio, 1160 feet per second single-stage fans, which was designed and tested as part of the NASA Experimental Quiet Engine Program. This fan was designed to deliver a bypass pressure ratio of 1.50 with an adiabatic efficiency of 86.5% at a total fan flow of 950 lb/sec. It was tested with and without inlet flow distortion. A bypass total-pressure ratio of 1.52 and an adiabatic efficiency of 88.3% at a total fan flow of 962 lb/sec were actually achieved. An operating margin of 12.4% was demonstrated at design speed.

Pre-Stall Behavior of a Transonic Axial Compressor Stage via Time-Accurate Numerical Simulation

NASA Technical Reports Server (NTRS)

Chen, Jen-Ping; Hathaway, Michael D.; Herrick, Gregory P.

2008-01-01

CFD calculations using high-performance parallel computing were conducted to simulate the pre-stall flow of a transonic compressor stage, NASA compressor Stage 35. The simulations were run with a full-annulus grid that models the 3D, viscous, unsteady blade row interaction without the need for an artificial inlet distortion to induce stall. The simulation demonstrates the development of the rotating stall from the growth of instabilities. Pressure-rise performance and pressure traces are compared with published experimental data before the study of flow evolution prior to the rotating stall. Spatial FFT analysis of the flow indicates a rotating long-length disturbance of one rotor circumference, which is followed by a spike-type breakdown. The analysis also links the long-length wave disturbance with the initiation of the spike inception. The spike instabilities occur when the trajectory of the tip clearance flow becomes perpendicular to the axial direction. When approaching stall, the passage shock changes from a single oblique shock to a dual-shock, which distorts the perpendicular trajectory of the tip clearance vortex but shows no evidence of flow separation that may contribute to stall.

Development of a Low-Reynolds Number, Nonlinear kappa-epsilon Model for the Reduced Navier-Stokes Equations

NASA Technical Reports Server (NTRS)

Boger, David A.; Govindan, T. R.; McDonald, Henry

1997-01-01

Previous work at NASA LeRC has shown that flow distortions in aircraft engine inlet ducts can be significantly reduced by mounting vortex generators, or small wing sections, on the inside surface of the engine inlet. The placement of the vortex generators is an important factor in obtaining the optimal effect over a wide operating envelope. In this regard, the only alternative to a long and expensive test program which would search out this optimal configuration is a good prediction procedure which could narrow the field of search. Such a procedure has been developed in collaboration with NASA LeRC, and results obtained by NASA personnel indicate that it shows considerable promise for predicting the viscous turbulent flow in engine inlet ducts in the presence of vortex generators. The prediction tool is a computer code which numerically solves the reduced Navier-Stokes equations and so is commonly referred to as RNS3D. Obvious deficiencies in RNS3D have been addressed in previous work. Primarily, it is known that the predictions of the mean velocity field of a turbulent boundary layer flow approaching separation are not in good agreement with data. It was suggested that the use of an algebraic mixing-length turbulence model in RNS3D is at least partly to blame for this. Additionally, the current turbulence model includes an assumption of isotropy which will ultimately fail to capture turbulence-driven secondary flow known to exist in noncircular ducts.

Development and Testing of a High Stability Engine Control (HISTEC) System

NASA Technical Reports Server (NTRS)

Orme, John S.; DeLaat, John C.; Southwick, Robert D.; Gallops, George W.; Doane, Paul M.

1998-01-01

Flight tests were recently completed to demonstrate an inlet-distortion-tolerant engine control system. These flight tests were part of NASA's High Stability Engine Control (HISTEC) program. The objective of the HISTEC program was to design, develop, and flight demonstrate an advanced integrated engine control system that uses measurement-based, real-time estimates of inlet airflow distortion to enhance engine stability. With improved stability and tolerance of inlet airflow distortion, future engine designs may benefit from a reduction in design stall-margin requirements and enhanced reliability, with a corresponding increase in performance and decrease in fuel consumption. This paper describes the HISTEC methodology, presents an aircraft test bed description (including HISTEC-specific modifications) and verification and validation ground tests. Additionally, flight test safety considerations, test plan and technique design and approach, and flight operations are addressed. Some illustrative results are presented to demonstrate the type of analysis and results produced from the flight test program.

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

NASA Astrophysics Data System (ADS)

Akturk, Ali

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

Flight Test Results of an Axisymmetric Channeled Center Body Supersonic Inlet at Off-Design Conditions

NASA Technical Reports Server (NTRS)

St. John, Clinton W.; Frederick, Michael Alan

2013-01-01

Flight-testing of a channeled center-body axisymmetric supersonic inlet design concept was conducted at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center in collaboration with the NASA Glenn Research Center (Cleveland, Ohio) and TechLand Research, Inc. (North Olmsted, Ohio). This testing utilized the Propulsion Flight Test Fixture, flown on the NASA F-15B research test bed airplane (NASA tail number 836) at local experiment Mach numbers up to 1.50. The translating channeled center-body inlet was designed by TechLand Research, Inc. (U.S. Patent No. 6,276,632 B1) to allow for a novel method of off-design flow matching, with original test planning conducted under a NASA Small Business Innovative Research study. Data were collected in flight at various off-design Mach numbers for fixed-geometry representations of both the channeled center-body design and an equivalent area smooth center-body design for direct comparison of total pressure recovery and limited distortion measurements.

Coupled parametric design of flow control and duct shape

NASA Technical Reports Server (NTRS)

Florea, Razvan (Inventor); Bertuccioli, Luca (Inventor)

2009-01-01

A method for designing gas turbine engine components using a coupled parametric analysis of part geometry and flow control is disclosed. Included are the steps of parametrically defining the geometry of the duct wall shape, parametrically defining one or more flow control actuators in the duct wall, measuring a plurality of performance parameters or metrics (e.g., flow characteristics) of the duct and comparing the results of the measurement with desired or target parameters, and selecting the optimal duct geometry and flow control for at least a portion of the duct, the selection process including evaluating the plurality of performance metrics in a pareto analysis. The use of this method in the design of inter-turbine transition ducts, serpentine ducts, inlets, diffusers, and similar components provides a design which reduces pressure losses and flow profile distortions.

Concept Designed and Developed for Distortion- Tolerant, High-Stability Engine Control

NASA Technical Reports Server (NTRS)

1995-01-01

Engine Control Future aircraft turbine engines, both commercial and military, must be able to successfully accommodate expected increased levels of steady-state and dynamic engine-face distortion. Advanced tactical aircraft are likely to use thrust vectoring to enhance their maneuverability. As a result, the engines will see more extreme aircraft angles-of-attack and sideslip levels than are currently encountered with present-day aircraft. Also, the mixed-compression inlets needed for the High Speed Civil Transport will likely encounter disturbances similar to those seen by tactical aircraft, in addition to planar pulse, inlet buzz, and high distortion levels at low flight speed and off-design operation. The current approach of incorporating a sufficient component design stall margin to tolerate these increased levels of distortion would significantly reduce performance. The objective of the High Stability Engine Control (HISTEC) program is to design, develop, and flight demonstrate an advanced, high-stability, integrated engine-control system that uses measurement-based, real-time estimates of distortion to enhance engine stability. The resulting distortion-tolerant control reduces the required design stall margin, with a corresponding increase in performance and decrease in fuel burn. The HISTEC concept has been designed and developed, and the software implementing the concept has successfully accommodated time-varying distortion. The NASA Lewis Research Center is currently overseeing the development and validation of the hardware and software necessary to flight test the HISTEC concept. HISTEC is a contracted effort with Pratt & Whitney of West Palm Beach, Florida. The HISTEC approach includes two major systems: A Distortion Estimation System (DES) and Stability Management Control (SMC). DES is an aircraft-mounted, high-speed processor that estimates the amount and type of distortion present and its effect on the engine. It uses high-response pressure measurements at the engine face to calculate indicators of the type and extent of distortion in real time. From these indicators, DES determines the effects of distortion on the propulsion systems and the corresponding engine match point necessary to accommodate it. DES output consists of fan and compressor pressure ratio trim commands that are passed to the SMC. In addition, DES uses maneuver information, consisting of angle-of-attack and sideslip from the flight control, to anticipate high inlet distortion conditions. The SMC, which is contained in the engine-mounted, Improved Digital Electronic Engine Control (IDEEC), includes advanced control laws to directly control the fan and compressor transient operating line (pressure ratio). These advanced control laws, with a multivariable design, have the potential for higher bandwidth and the resulting more precise control of engine match. The ability to measure and assess the distortion effects in real time coupled with a high-response controller improves engine stability at high levels of distortion. The software algorithms implementing DES have been designed, developed, and demonstrated, and integration testing of the DES and SMC software has been completed. The results show that the HISTEC system will be able to sense inlet distortion, determine the effect on engine stability, and accommodate distortion by maintaining an adequate margin for engine surge. The Pratt &Whitney Comprehensive Engine Diagnostic Unit was chosen as the DES processor. An instrumented inlet case for sensing distortion was designed and fabricated. HISTEC is scheduled for flight test on the ACTIVE F-15 aircraft at the NASA Dryden Flight Research Center in Edwards, California, in late 1996.

Hot gas ingestion test results of a two-poster vectored thrust concept with flow visualization in the NASA Lewis 9- by 15-foot low speed wind tunnel

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Neiner, George; Bencic, Timothy J.; Flood, Joseph D.; Amuedo, Kurt C.

1990-01-01

A 9.2 percent scale STOVL hot gas ingestion model was tested in the NASA Lewis 9 x 15-foot Low-Speed Wind Tunnel. Flow visualization from the Phase 1 test program, which evaluated the hot ingestion phenomena and control techniques, is covered. The Phase 2 test program evaluated the hot gas ingestion phenomena at higher temperatures and used a laser sheet to investigate the flow field. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/life improvement devices (LIDs) which reduced the hot gas ingestion. The test was conducted at full scale nozzle pressure ratios and inlet Mach numbers. Results are presented over a range of nozzle pressure ratios at a 10 kn headwind velocity. The Phase 2 program was conducted at exhaust nozzle temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. The results reported are for nozzle exhaust temperatures up to 1160 R and contain the compressor face pressure and temperature distortions, the total pressure recovery, the inlet temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane contours, the model airframe heating, and the location of the ground flow separation.

A 727 airplane center duct inlet low speed performance confirmation model test for refanned JT8D engines, phase 2

NASA Technical Reports Server (NTRS)

Kaldschmidt, G.; Syltebo, B. E.; Ting, C. T.

1973-01-01

The results from testing of a 0.3 scale model center duct inlet (S duct) for the Pratt and Whitney Aircraft JT8D-100 engines are presented. The objective of this test was to demonstrate that the required airflow of the JT8D-100 engine (480 lb/sec as compared to 334 lb/sec for JT8D-15) can be achieved with minimum modifications to the existing 727 airplane structure at acceptable levels of total pressure recovery and distortion. Steady-state pressure recovery, steady-state pressure distortion, and dynamic pressure measurements were taken at the engine face station. Surface static pressure measurements were taken along the duct. Test results indicated that the required airflow was achieved with acceptable pressure recovery (comparable to the current 727-200 S duct). Inlet inflow angle variation within the 727 airplane operating regime (minus 5 to 5 degrees) had no effect on the inlet performance. Pressure distortion at static and forward speed at takeoff airflow conditions are within P and WA limits for the Phase II duct when equipped with vortex generators. Static crosswind operation between 10 knots and 25 knots appears feasible at full takeoff power.

Performance of a Blunt-lip Side Inlet with Ramp Bleed, Bypass, and a Long Constant-area Duct Ahead of the Engine : Mach Numbers 0.66 and 1.5 to 2.1

NASA Technical Reports Server (NTRS)

Allen, John L

1956-01-01

Unsteady shock-induced separation of the ramp boundary layer was reduced and stabilized more effectively by external perforations than by external or internal slots. At Mach 2.0 peak total-pressure recovery was increased from 0.802 to 0.89 and stable mass-flow range was increased 185 percent over that for the solid ramp. Peak pressure recovery occurred just before instability. The 7 and one-third-diameter duct ahead of the engine reduced large total-pressure distortions but was not as successful for small distortions as obtained with throat bleed. By removing boundary-layer air the bypass nearly recovered the total-pressure loss due to the long duct.

Factors Affecting Inlet-Engine Compatibility During Aircraft Departures at High Angle of Attack for an F/A -18A Aircraft

NASA Technical Reports Server (NTRS)

Steenken, W. G.; Williams, J. G.; Yuhas, A. J.; Walsh, K. R.

1999-01-01

The F404-GE-400 engine 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. A series of 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 yaw rates at departure recovery were in the range of 40 to 90 degrees per second. 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 were determined to not significantly exceed those measured at maximum angle-of-attack and - sideslip maneuvers during controlled flight. Surges as a result of inlet distortion levels were anticipated to initiate in the fan. Analysis revealed that the surges initiated in the compressor and were the result of a combination of high levels of inlet distortion and rapid changes in aircraft motion. These rapid changes in aircraft motion are indicative of a combination of engine mount 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.

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

NASA Astrophysics Data System (ADS)

Herges, Thomas G.

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

Data reduction formulas for the 16-foot transonic tunnel: NASA Langley Research Center, revision 2

NASA Technical Reports Server (NTRS)

Mercer, Charles E.; Berrier, Bobby L.; Capone, Francis J.; Grayston, Alan M.

1992-01-01

The equations used by the 16-Foot Transonic Wind Tunnel in the data reduction programs are presented in nine modules. Each module consists of equations necessary to achieve a specific purpose. These modules are categorized in the following groups: (1) tunnel parameters; (2) jet exhaust measurements; (3) skin friction drag; (4) balance loads and model attitudes calculations; (5) internal drag (or exit-flow distribution); (6) pressure coefficients and integrated forces; (7) thrust removal options; (8) turboprop options; and (9) inlet distortion.

Turbine endwall two-cylinder program. [wind tunnel and water tunnel investigation of three dimensional separation of fluid flow

NASA Technical Reports Server (NTRS)

Langston, L. S.

1980-01-01

Progress is reported in an effort to study the three dimensional separation of fluid flow around two isolated cylinders mounted on an endwall. The design and performance of a hydrogen bubble generator for water tunnel tests to determine bulk flow properties and to measure main stream velocity and boundary layer thickness are described. Although the water tunnel tests are behind schedule because of inlet distortion problems, tests are far enough along to indicate cylinder spacing, wall effects and low Reynolds number behavior, all of which impacted wind tunnel model design. The construction, assembly, and operation of the wind tunnel and the check out of its characteristics are described. An off-body potential flow program was adapted to calculate normal streams streamwise pressure gradients at the saddle point locations.

The Performance of a Subsonic Diffuser Designed for High Speed Turbojet-Propelled Flight

NASA Technical Reports Server (NTRS)

Biesiadny, Thomas J. (Technical Monitor); Wendt, Bruce J.

2004-01-01

An initial-phase subsonic diffuser has been designed for the turbojet flowpath of the hypersonic x43B flight demonstrator vehicle. The diffuser fit into a proposed mixed-compression supersonic inlet system and featured a cross-sectional shape transitioning flowpath (high aspect ratio rectangular throat-to-circular engine face) and a centerline offset. This subsonic diffuser has been fabricated and tested at the W1B internal flow facility at NASA Glenn Research Center. At an operating throat Mach number of 0.79, baseline Pitot pressure recovery was found to be just under 0.9, and DH distortion intensity was about 0.4 percent. The diffuser internal flow stagnated, but did not separate on the offset surface of this initial-phase subsonic diffuser. Small improvements in recovery (+0.4 percent) and DH distortion (-32 percent) were obtained from using vane vortex generator flow control applied just downstream of the diffuser throat. The optimum vortex generator array patterns produced inflow boundary layer divergence (local downwash) on the offset surface centerline of the diffuser, and an inflow boundary layer convergence (local upwash) on the centerline of the opposite surface.

An Experimental Study of Fan Inflow Distortion Tone Noise

NASA Technical Reports Server (NTRS)

Koch, L. Danielle

2010-01-01

The tone noise generated when a fan ingests circumferentially distorted flow was studied by an experiment conducted with the Advanced Noise Control Fan at the NASA Glenn Research Center. The inflow was distorted by inserting cylindrical rods radially into the duct. The rods were arranged in circumferentially irregular patterns in three of the five configurations tested. Rods were held in place using a mounting ring with 30 equally spaced holes placed at an axial location one rotor chordlength upstream of the fan. Acoustic pressure was measured in the inlet and exhaust duct of the fan using the Rotating Rake fan tone measurement system. Sound power levels, calculated from the measured data, were plotted as a function of circumferential mode. An analytic description of the unsteady pressure distribution at the interaction plane between the stationary rods and the fan rotor is presented in a form suitable for representing the circumferentially irregularly placed rods. Terms in the analytical description for sound power were proven to be useful in determining the dominant circumferential modes measured in the experiment and the differences in mode power level between the configurations tested. Insight gained through this work will be useful in the development of tools to compute fan inflow distortion tone noise.

Performance of a multiple venturi fuel-air preparation system. [fuel injection for gas turbines

NASA Technical Reports Server (NTRS)

Tacina, R. R.

1979-01-01

Spatial fuel-air distributions, degree of vaporization, and pressure drop were measured 16.5 cm downstream of the fuel injection plane of a multiple Venturi tube fuel injector. Tests were performed in a 12 cm tubular duct. Test conditions were: a pressure of 0.3 MPa, inlet air temperature from 400 to 800K, air velocities of 10 and 20 m/s, and fuel-air ratios of 0.010 and 0.020. The fuel was Diesel #2. Spatial fuel-air distributions were within + or - 20 percent of the mean at inlet air temperatures above 450K. At an inlet air temperature of 400K, the fuel-air distribution was measured when a 50 percent blockage plate was placed 9.2 cm upstream of the fuel injection plane to distort the inlet air velocity fuel injection plane to distort the inlet air velocity profile. Vaporization of the fuel was 50 percent complete at an inlet air temperature of 400K and the percentage increased linearly with temperature to complete vaporization at 600K. The pressure drop was 3 percent at the design point which was three times greater than the designed value and the single tube experiment value. No autoignition or flashback was observed at the conditions tested.

Experimental quiet engine program aerodynamic performance of Fan B

NASA Technical Reports Server (NTRS)

Giffin, R. G.; Parker, D. E.; Dunbar, L. W.

1972-01-01

This report presents the aerodynamic component test results of Fan B, one of two high-bypass-ratio, 1160 feet per second (353.6 m/sec) single-stage fans, which was designed and tested as part of the NASA Experimental Quiet Engine Program. The fan was designed to deliver a bypass pressure ratio of 1.50 with an adiabatic efficiency of 87.0% at a total fan flow of 950 lb/sec (430.9 kg/sec). It was tested with and without inlet distortion. A bypass total pressure ratio of 1.52 and an adiabatic efficiency of 86.9% at a total fan flow of 966 lb/sec (438.2 kg/sec) were actually achieved. An operating margin of 19.5% was demonstrated at design speed.

Tidal Wave Reflectance, Evolution and Distortion in Elkhorn Slough, CA

DTIC Science & Technology

2013-03-01

School O1 Lunisolar diurnal Tidal Constituent ONR Office of Naval Research p Pressure Rhfm High-Frequency Motion Tidal Reflection Coefficient RIVET ...2012 an experiment at the New River Inlet, known as the River and Inlet Dynamics experiment ( RIVET ) was conducted. RIVET 2 is currently scheduled for

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

AeroPropulsoServoElasticity: Dynamic Modeling of the Variable Cycle Propulsion System

NASA Technical Reports Server (NTRS)

Kopasakis, George

2012-01-01

This presentation was made at the 2012 Fundamental Aeronautics Program Technical Conference and it covers research work for the Dynamic Modeling of the Variable cycle Propulsion System that was done under the Supersonics Project, in the area of AeroPropulsoServoElasticity. The presentation covers the objective for the propulsion system dynamic modeling work, followed by the work that has been done so far to model the variable Cycle Engine, modeling of the inlet, the nozzle, the modeling that has been done to model the affects of flow distortion, and finally presenting some concluding remarks and future plans.

HIMAT Inlet Model in the 8- by 6-Foot Supersonic Wind Tunnel

NASA Image and Video Library

1979-02-21

A Highly Maneuverable Aircraft Technology (HiMAT) inlet model installed in the test section of the 8- by 6-Foot Supersonic Wind Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Engineers at the Ames Research Center, Dryden Flight Research Center, and Rockwell International designed two pilotless subscale HiMAT vehicles in the mid-1970s to study new design concepts for fighter aircraft in the transonic realm without risking the lives of test pilots. The aircraft used sophisticated technologies such as advanced aerodynamics, composite materials, digital integrated propulsion control, and digital fly-by-wire control systems. In late 1977 NASA Lewis studied the HiMAT’s General Electric J85-21 jet engine in the Propulsion Systems Laboratory. The researchers charted the inlet quality with various combinations anti-distortion screens. HiMAT employed a relatively short and curved inlet compared to actual fighter jets. In the spring of 1979, Larry Smith led an in-depth analysis of the HiMAT inlet in the 8- by 6 tunnel. The researchers installed vortex generators to battle flow separation in the diffuser. The two HiMAT aircraft performed 11 hours of flying over the course of 26 missions from mid-1979 to January 1983 at Dryden and Ames. Although the HiMAT vehicles were considered to be overly complex and expensive, the program yielded a wealth of data that would validate computer-based design tools.

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

NASA Technical Reports Server (NTRS)

Boles, Michael A.; Heavner, Richard L.

1988-01-01

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

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.

Task II: Three-dimensional Rotating Stall Inception and Effects of Rotating Tip Clearance Asymmetry in Axial Compressors

NASA Technical Reports Server (NTRS)

Suder, Kenneth (Technical Monitor); Tan, Choon-Sooi

2003-01-01

The effects of two types of flow non-uniformity on stall inception behavior were assessed with linearized stability analyses of two compressor flow models. Response to rotating tip clearance asymmetries induced by a whirling rotor shaft or rotor height variations were investigated with a two-dimensional flow model. A 3-D compressor model was also developed to study the stability of both full-span and part-span rotating stall modes in annular geometries with radial flow variations. The studies focussed on (1) understanding what compressor designs were sensitive to these types of circumferential and spanwise flow non-uniformities, and (2) situations where 2-D stability theories were inadequate because of 3-D flow effects. Rotating tip clearance non-uniformity caused the greatest performance loss for shafts whirling at the rotating stall frequency. A whirling shaft displacement of 1 percent chord caused the stalling mass flow to rise by as much as 10 percent and the peak pressure rise to decrease by 6 percent. These changes were an order of magnitude larger than for equivalent-sized stationary or rotor-locked clearance asymmetries. Spanwise flow non-uniformity always destabilized the compressor, so that 2-D models over-predicted that stall margin compared to 3-D theory. The difference increased for compressors with larger spanwise variations of characteristic slope and reduced characteristic curvature near the peak. Differences between 2-D and 3-D stall point predictions were generally unacceptable (2 - 4 percent of flow coefficient) for single-stage configurations, but were less than 1 percent for multistage compressors. 2-D analyses predicted the wrong stall mode for specific cases of radial inlet flow distortion, mismatching and annulus area contraction, where higher-order radial modes led to stall. The stability behavior of flows with circumferential or radial non-uniformity was unified through a single stability criterion. The stall point for both cases was set by the integral around the annulus of the pressure rise characteristic slope, weighted by the amplitude of the mode shape. For the case of steady circumferential variations, this criterion reduced to the integrated mean slope (IMS) condition associated with steady inlet distortions. The rotating tip clearance asymmetry model was also used to demonstrate the feasibility of actively controlling the shaft position to suppress rotating stall. In axisymmetric mean flow, this method only stabilized the first harmonic mode, increasing the operating range until surge or higher harmonic modes became unstable.

Predicted and Measured Modal Sound Power Levels for a Fan Ingesting Distorted Inflow

NASA Technical Reports Server (NTRS)

Koch, L. Danielle

2010-01-01

Refinements have been made to a method for estimating the modal sound power levels of a ducted fan ingesting distorted inflow. By assuming that each propagating circumferential mode consists only of a single radial mode (the one with the highest cut-off ratio), circumferential mode sound power levels can be computed for a variety of inflow distortion patterns and operating speeds. Predictions from the refined theory have been compared to data from an experiment conducted in the Advanced Noise Control Fan at NASA Glenn Research Center. The inflow to the fan was distorted by inserting cylindrical rods radially into the inlet duct. The rods were placed at an axial location one rotor chord length upstream of the fan and arranged in both regular and irregular circumferential patterns. The fan was operated at 2000, 1800, and 1400 rpm. Acoustic pressure levels were measured in the fan inlet and exhaust ducts using the Rotating Rake fan mode measurement system. Far field sound pressure levels were also measured. It is shown that predicted trends in circumferential mode sound power levels closely match the experimental data for all operating speeds and distortion configurations tested. Insight gained through this work is being used to develop more advanced tools for predicting fan inflow distortion tone noise levels.

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.

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.

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

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

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

1999-07-01

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

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.

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.

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-07-01

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

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.

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

NASA Technical Reports Server (NTRS)

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

1995-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1976-01-01

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

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.

Basic research in fan source noise: Inlet distortion and turbulence noise

NASA Technical Reports Server (NTRS)

Kantola, R. A.; Warren, R. E.

1978-01-01

A widely recognized problem in jet engine fan noise is the discrepancy between inflight and static tests. This discrepancy consists of blade passing frequency tones, caused by ingested turbulence that appear in the static tests but not in flight. To reduce the ingested distortions and turbulence in an anechoic chamber, a reverse cone inlet is used to guide the air into the fan. This inlet also has provisions for boundary layer suction and is used in conjunction with a turbulence control structure (TCS) to condition the air impinging on the fan. The program was very successful in reducing the ingested turbulence, to the point where reductions in the acoustic power at blade passing frequency are as high as 18 db for subsonic tip speeds. Even with this large subsonic tone suppression, the supersonic tip speed tonal content remains largely unchanged, indicating that the TCS did not appreciably attenuate the noise but effects the generation via turbulence reduction. Turbulence mapping of the inlet confirmed that the tone reductions are due to a reduction in turbulence, as the low frequency power spectra of the streamwise and transverse turbulence were reduced by up to ten times and 100 times, respectively.

Three-Dimensional Aerodynamic Instabilities In Multi-Stage Axial Compressors

NASA Technical Reports Server (NTRS)

Tan, Choon S.; Gong, Yifang; Suder, Kenneth L. (Technical Monitor)

2001-01-01

This thesis presents the conceptualization and development of a computational model for describing three-dimensional non-linear disturbances associated with instability and inlet distortion in multistage compressors. Specifically, the model is aimed at simulating the non-linear aspects of short wavelength stall inception, part span stall cells, and compressor response to three-dimensional inlet distortions. The computed results demonstrated the first-of-a-kind capability for simulating short wavelength stall inception in multistage compressors. The adequacy of the model is demonstrated by its application to reproduce the following phenomena: (1) response of a compressor to a square-wave total pressure inlet distortion; (2) behavior of long wavelength small amplitude disturbances in compressors; (3) short wavelength stall inception in a multistage compressor and the occurrence of rotating stall inception on the negatively sloped portion of the compressor characteristic; (4) progressive stalling behavior in the first stage in a mismatched multistage compressor; (5) change of stall inception type (from modal to spike and vice versa) due to IGV stagger angle variation, and "unique rotor tip incidence" at these points where the compressor stalls through short wavelength disturbances. The model has been applied to determine the parametric dependence of instability inception behavior in terms of amplitude and spatial distribution of initial disturbance, and intra-blade-row gaps. It is found that reducing the inter-blade row gaps suppresses the growth of short wavelength disturbances. It is also concluded from these parametric investigations that each local component group (rotor and its two adjacent stators) has its own instability point (i.e. conditions at which disturbances are sustained) for short wavelength disturbances, with the instability point for the compressor set by the most unstable component group. For completeness, the methodology has been extended to describe finite amplitude disturbances in high-speed compressors. Results are presented for the response of a transonic compressor subjected to inlet distortions.

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.

Design and development of an F/A-18 inlet distortion rake: A cost and time saving solution

NASA Technical Reports Server (NTRS)

Yuhas, Andrew J.; Ray, Ronald J.; Burley, Richard R.; Steenken, William G.; Lechtenberg, Leon; Thornton, Don

1995-01-01

An innovative inlet total pressure distortion measurement rake has been designed and developed for the F/A-18 A/B/C/D aircraft inlet. The design was conceived by NASA and General Electric Aircraft Engines personnel. This rake has been flight qualified and flown in the F/A-18 High Alpha Research Vehicle at NASA Dryden Flight Research Center, Edwards, California. The eight-legged, one-piece, wagon wheel design of the rake was developed at a reduced cost and offered reduced installation time compared to traditional designs. The rake features 40 dual-measurement ports for low- and high-frequency pressure measurements with the high-frequency transducer mounted at the port. This high-frequency transducer offers direct absolute pressure measurements from low to high frequencies of interest, thereby allowing the rake to be used during highly dynamic aircraft maneuvers. Outstanding structural characteristics are inherent to the design through its construction and use of lightweight materials.

Aeromechanics Analysis of a Distortion-Tolerant Fan with Boundary Layer Ingestion

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Reddy, T. S. R.; Coroneos, Rula M.; Min, James B.; Provenza, Andrew J.; Duffy, Kirsten P.; Stefko, George L.; Heinlein, Gregory S.

2018-01-01

A propulsion system with Boundary Layer Ingestion (BLI) has the potential to significantly reduce aircraft engine fuel burn. But a critical challenge is to design a fan that can operate continuously with a persistent BLI distortion without aeromechanical failure -- flutter or high cycle fatigue due to forced response. High-fidelity computational aeromechanics analysis can be very valuable to support the design of a fan that has satisfactory aeromechanic characteristics and good aerodynamic performance and operability. Detailed aeromechanics analyses together with careful monitoring of the test article is necessary to avoid unexpected problems or failures during testing. In the present work, an aeromechanics analysis based on a three-dimensional, time-accurate, Reynolds-averaged Navier-Stokes computational fluid dynamics code is used to study the performance and aeromechanical characteristics of the fan in both circumferentially-uniform and circumferentially-varying distorted flows. Pre-test aeromechanics analyses are used to prepare for the wind tunnel test and comparisons are made with measured blade vibration data after the test. The analysis shows that the fan has low levels of aerodynamic damping at various operating conditions examined. In the test, the fan remained free of flutter except at one near-stall operating condition. Analysis could not be performed at this low mass flow rate operating condition since it fell beyond the limit of numerical stability of the analysis code. The measured resonant forced response at a specific low-response crossing indicated that the analysis under-predicted this response and work is in progress to understand possible sources of differences and to analyze other larger resonant responses. Follow-on work is also planned with a coupled inlet-fan aeromechanics analysis that will more accurately represent the interactions between the fan and BLI distortion.

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.

Aeromechanics Analysis of a Boundary Layer Ingesting Fan

NASA Technical Reports Server (NTRS)

Bakhle, Milind A.; Reddy, T. S. R.; Herrick, Gregory P.; Shabbir, Aamir; Florea, Razvan V.

2013-01-01

Boundary layer ingesting propulsion systems have the potential to significantly reduce fuel burn but these systems must overcome the challe nges related to aeromechanics-fan flutter stability and forced response dynamic stresses. High-fidelity computational analysis of the fan a eromechanics is integral to the ongoing effort to design a boundary layer ingesting inlet and fan for fabrication and wind-tunnel test. A t hree-dimensional, time-accurate, Reynolds-averaged Navier Stokes computational fluid dynamics code is used to study aerothermodynamic and a eromechanical behavior of the fan in response to both clean and distorted inflows. The computational aeromechanics analyses performed in th is study show an intermediate design iteration of the fan to be flutter-free at the design conditions analyzed with both clean and distorte d in-flows. Dynamic stresses from forced response have been calculated for the design rotational speed. Additional work is ongoing to expan d the analyses to off-design conditions, and for on-resonance conditions.

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.

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.

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

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Weir, Lois

2014-01-01

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

Computational effects of inlet representation on powered hypersonic, airbreathing models

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-03-01

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

An experimental evaluation of S-duct inlet-diffuser configurations for turboprop offset gearbox applications

NASA Technical Reports Server (NTRS)

Mcdill, Paul L.

1986-01-01

A test program, utilizing a large scale model, was run in the NASA Lewis Research Center 10- by 10-ft wind tunnel to examine the influence on performance of design parameters of turboprop S-duct inlet/diffuser systems. The parametric test program investigated inlet lip thickness, inlet/diffuser cross-sectional geometry, throat design Mach number, and shaft fairing shape. The test program was run at angles of attack to 15 deg and tunnel Mach numbers to 0.35. Results of the program indicate that current design techniques can be used to design inlet/diffuser systems with acceptable total pressure recovery, but several of the design parameters, notably lip thickness (contraction ratio) and shaft fairing cross section, must be optimized to prevent excessive distortion at the compressor face.

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.

Flow and Performance Calculations of Axial Compressor near Stall Margin

NASA Astrophysics Data System (ADS)

Hwang, Yoojun; Kang, Shin-Hyoung

2010-06-01

Three-dimensional flows through a Low Speed Research Axial Compressor were numerically conducted in order to estimate the performance through unsteady and steady-state simulations. The first stage with the inlet guide vane was investigated at the design point to confirm that the rotor blade induced periodicity exists. Special attention was paid to the flow near the stall condition to inspect the flow behavior in the vicinity of the stall margin. The performance predicted under the steady-state assumption is in good agreement with the measured data. However, the steady-state calculations induce more blockage through the blade passage. Flow separations on the blade surface and end-walls are reduced when unsteady simulation is conducted. The negative jet due to the wake of the rotor blade periodically distorts the boundary layer on the surface of the stator blade and improves the performance of the compressor in terms of the pressure rise. The advantage of the unsteadiness increases as the flow rate reduces. In addition, the rotor tip leakage flow is forced downstream by the unsteadiness. Consequently, the behavior contributes to extending the range of operation by preventing the leakage flow from proceeding upstream near the stall margin.

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

NASA Technical Reports Server (NTRS)

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

1979-01-01

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

Three dimensional viscous analysis of a hypersonic inlet

NASA Technical Reports Server (NTRS)

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

1989-01-01

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

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.

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.

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

NASA Technical Reports Server (NTRS)

Benson, Thomas J.

1994-01-01

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

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

NASA Astrophysics Data System (ADS)

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

1991-02-01

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

Effect of Pulsed Plasma Jets on the Recovering Boundary Layer Downstream of a Reflected Shock Interaction

NASA Astrophysics Data System (ADS)

Greene, Benton; Clemens, Noel; Magari, Patrick; Micka, Daniel; Ueckermann, Mattheus

2015-11-01

Shock-induced turbulent boundary layer separation can have many detrimental effects in supersonic inlets including flow distortion and instability, structural fatigue, poor pressure recovery, and unstart. The current study investigates the effect of pulsed plasma jets on the recovering boundary layer downstream of a reflected shock wave-boundary layer interaction. The effects of pitch and skew angle of the jet as well as the heating parameter and discharge time scale are tested using several pulsing frequencies. In addition, the effect of the plasma jets on the undisturbed boundary layer at 6 mm and 11 mm downstream of the jets is measured. A pitot-static pressure probe is used to measure the velocity profile of the boundary layer 35 mm downstream of the plasma jets, and the degree of boundary layer distortion is compared between the different models and run conditions. Additionally, the effect of each actuator configuration on the shape of the mean separated region is investigated using surface oil flow visualization. Previous studies with lower energy showed a weak effect on the downstream boundary layer. The current investigation will attempt to increase this effect using a higher-energy discharge. Funded by AFRL through and SBIR in collaboration with Creare, LLC.

Performance Calculations for a Boundary-Layer-Ingesting Fan Stage from Sparse Measurements

NASA Technical Reports Server (NTRS)

Hirt, Stefanie M.; Wolter, John D.; Arend, David J.; Hearn, Tristan A.; Hardin, Larry W.; Gazzaniga, John A.

2018-01-01

A test of the Boundary Layer Ingesting-Inlet / Distortion-Tolerant Fan was completed in NASA Glenn's 8-Foot by 6-Foot supersonic wind tunnel. Inlet and fan performance were measured by surveys using a set of rotating rake arrays upstream and downstream of the fan stage. Surveys were conducted along the 100 percent speed line and a constant exit corrected flow line passing through the aerodynamic design point. These surveys represented only a small fraction of the data collected during the test. For other operating points, data was recorded as snapshots without rotating the rakes which resulted in a sparser set of recorded data. This paper will discuss analysis of these additional, lower measurement density data points to expand our coverage of the fan map. Several techniques will be used to supplement the snapshot data at test conditions where survey data also exists. The supplemented snapshot data will be compared with survey results to assess the quality of the approach. Effective methods will be used to analyze the data set for which only snapshots exist.

Assessment of Integrated Nozzle Performance

NASA Technical Reports Server (NTRS)

Lambert, H. H.; Mizukami, M.

1999-01-01

This presentation highlights the activities that researchers at the NASA Lewis Research Center (LeRC) have been and will be involved in to assess integrated nozzle performance. Three different test activities are discussed. First, the results of the Propulsion Airframe Integration for High Speed Research 1 (PAIHSR1) study are presented. The PAIHSR1 experiment was conducted in the LeRC 9 ft x l5 ft wind tunnel from December 1991 to January 1992. Second, an overview of the proposed Mixer/ejector Inlet Distortion Study (MIDIS-E) is presented. The objective of MIDIS-E is to assess the effects of applying discrete disturbances to the ejector inlet flow on the acoustic and aero-performance of a mixer/ejector nozzle. Finally, an overview of the High-Lift Engine Aero-acoustic Technology (HEAT) test is presented. The HEAT test is a cooperative effort between the propulsion system and high-lift device research communities to assess wing/nozzle integration effects. The experiment is scheduled for FY94 in the NASA Ames Research Center (ARC) 40 ft x 80 ft Low Speed Wind Tunnel (LSWT).

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

NASA Technical Reports Server (NTRS)

Stockman, N. O.

1975-01-01

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

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

NASA Technical Reports Server (NTRS)

Dittrich, R. T.

1972-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-03-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-01-01

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

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

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.

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.

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.

Flow Characterization and Dynamic Analysis of a Radial Compressor with Passive Method of Surge Control

NASA Astrophysics Data System (ADS)

Guillou, Erwann

Due to recent emission regulations, the use of turbochargers for force induction of internal combustion engines has increased. Actually, the trend in diesel engines is to downsize the engine by use of turbochargers that operate at higher pressure ratio. Unfortunately, increasing the rotational speed tends to reduce the turbocharger radial compressor range of operation which is limited at low mass flow rate by the occurrence of surge. In order to extent the operability of turbochargers, compressor housings can be equipped with a passive surge control device also known as ported shroud. This specific casing treatment has been demonstrated to enhance surge margin with minor negative impact on the compressor efficiency. However, the actual working mechanisms of the bypass system remain not well understood. In order to optimize the design of the ported shroud, it is then crucial to identify the dynamic flow changes induced by the implementation of the device to control instabilities. Experimental methods were used to assess the development of instabilities from stable, stall and eventually surge regimes of a ported shroud centrifugal compressor. Systematic comparison was conducted with the same compressor design without ported shroud. Hence, the full pressure dynamic survey of both compressors' performance characteristics converged toward two different and probably interrelated driving mechanisms to the development and/or propagation of unsteadiness within each compressor. One related the pressure disturbances at the compressor inlet, and notably the more apparent development of perturbations in the non-ported compressor impeller, whereas the other was attributed to the pressure distortions induced by the presence of the tongue in the asymmetric design of the compressor volute. Specific points of operation were selected to carry out planar flow measurements. At normal working, both standard and stereoscopic particle imaging velocimetry (PIV) measurements were performed to calculate the instantaneous and mean velocity fields at the inlet of the compressor. At incipient and full surge, phase-locked PIV measurements were added. In this work, satisfying characterization of the compressor inlet flow instabilities was obtained at different operational speeds. Combining transient pressure data and PIV measurements, the time evolution of the complex flow patterns occurring at surge was reconstructed and a better insight into the bypass mechanisms was achieved.

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.

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

NASA Technical Reports Server (NTRS)

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

1952-01-01

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

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.

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

Engine Response to Distorted Inflow Conditions: Conference Proceedings of the Propulsion and Energetics Specialists’ Meeting (68th) Held in Munich, Germany on 8-9 September 1986.

DTIC Science & Technology

1987-03-01

DISTORTION UNSTEADY INLET DISTORTION CHARACTERISTICS WITH THE B-1 B by C.J.MacMiler and W.R.Haagenson 16 DETERMINATION EXPERIMENTALE DES LOIS DE TRANSFERT DE...compresseurs. Dens ce domaine. los trois principaux objectifs du motoriste sont los suivants (a) - d6mont rer avant los essais en vol. la compatibilitA...ann~es sur la conception des moteurs deavions. 2 - ETHODES DE CALCUL 2.1. METNODES DES SECTEURS DE COMPRESSEUR EN PARALLELE 2.1.1. Description do la m

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.

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).

Investigation of Injector Slot Geometry on Curved-Diffuser Aerodynamic Performance

NASA Technical Reports Server (NTRS)

Silva, Odlanier

2004-01-01

The Compressor Branch vision is to be recognized as world-class leaders in research for fluid mechanics of compressors. Its mission is to conduct research and develop technology to advance the state of the art of compressors and transfer new technology to U.S. industries. Maintain partnerships with U.S. industries, universities, and other government organizations. Maintain a balance between customers focused and long range research. Flow control comprises enabling technologies to meet compression system performance requirements driven by emissions and fuel reduction goals (e.g., in UEET), missions (e.g., access-to-space), aerodynamically aggressive vehicle configurations (e.g., UAV and future blended wing body configurations with highly distorted inlets), and cost goals (e.g., in VAATE). The compression system requirements include increased efficiency, power-to-weight, and adaptability (i.e., robustness in terms of wide operability, distortion tolerance, and engine system health and reliability). The compressor flow control task comprises efforts to develop, demonstrate, and transfer adaptive flow control technology to industry to increase aerodynamic loading at current blade row loss levels, to enable adaptive1 y wide operability, and to develop plant models for adaptive compression systems. In this context, flow control is the controlled modification of a flow field by a deliberate means beyond the natural (uncontrolled) shaping of the solid surfaces that define the principal flow path. The objective of the compressor flow control task is to develop and apply techniques that control circulation, aerodynamic blockage, and entropy production in order to enhance the performance and operability of compression systems for advanced aero-propulsion applications. This summer I would be working with a curved-diffuser because it simulates what happens with flow in the stator blades in the compressor. With this experiment I will be doing some data analysis and parametric study of the injector slot geometries to get the best aerodynamic performance of it. This includes some data reduction, redesign and fast prototyping of the injector nozzle.

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.

Three-dimensional analysis of the Pratt and Whitney alternate design SSME fuel turbine

NASA Technical Reports Server (NTRS)

Kirtley, K. R.; Beach, T. A.; Adamczyk, J. J.

1991-01-01

The three dimensional viscous time-mean flow in the Pratt and Whitney alternate design space shuttle main engine fuel turbine is simulated using the average passage Navier-Stokes equations. The migration of secondary flows generated by upstream blade rows and their effect on the performance of downstream blade rows is studied. The present simulation confirms that the flow in this two stage turbine is highly three dimensional and dominated by the tip leakage flow. The tip leakage vortex generated by the first blade persists through the second blade and adversely affects its performance. The greatest mixing of the inlet total temperature distortion occurs in the second vane and is due to the large leakage vortex generated by the upstream rotor. It is assumed that the predominant spanwise mixing mechanism in this low aspect ratio turbine is the radial transport due to the deterministically unsteady vortical flow generated by upstream blade rows. A by-product of the analysis is accurate pressure and heat loads for all blade rows under the influence of neighboring blade rows. These aero loads are useful for advanced structural analysis of the vanes and blades.

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

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.

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

Advanced two-stage compressor program design of inlet stage

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

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

NASA Technical Reports Server (NTRS)

Bozak, Richard F.

2018-01-01

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

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

NASA Astrophysics Data System (ADS)

Kulkarni, P. S.

2017-04-01

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

Multipurpose Rotating Rake Arrays for Integrated Inlet and Fan Stage Performance Measurement

NASA Technical Reports Server (NTRS)

Wolter, John D.

2017-01-01

Low-pressure-ratio fan stage performance measurement requires precise measurement of conditions upstream and downstream of the fan stage. This presentation will discuss the rotating rake arrays used for the recent Boundary Layer Ingesting Inlet-Distortion-Tolerant Fan experiment in the NASA Glenn 8 by 6 foot wind tunnel. To achieve precise measurements, simulations of the rake sampling from pre-test CFD (Computerized Fluid Dynamics) solutions were used to optimize the number and locations of rake instruments.

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.

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.

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

NASA Astrophysics Data System (ADS)

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

2016-04-01

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

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

NASA Technical Reports Server (NTRS)

Slater, John W.

2002-01-01

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

Harmonic uniflow engine

DOEpatents

Bennett, Charles L.

2016-03-22

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

Turboprop engine and method of operating the same

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

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

1986-02-11

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

Estimation of additive forces and moments for supersonic inlets

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

A Three-Dimensional Unsteady CFD Model of Compressor Stability

NASA Technical Reports Server (NTRS)

Chima, Rodrick V.

2006-01-01

A three-dimensional unsteady CFD code called CSTALL has been developed and used to investigate compressor stability. The code solved the Euler equations through the entire annulus and all blade rows. Blade row turning, losses, and deviation were modeled using body force terms which required input data at stations between blade rows. The input data was calculated using a separate Navier-Stokes turbomachinery analysis code run at one operating point near stall, and was scaled to other operating points using overall characteristic maps. No information about the stalled characteristic was used. CSTALL was run in a 2-D throughflow mode for very fast calculations of operating maps and estimation of stall points. Calculated pressure ratio characteristics for NASA stage 35 agreed well with experimental data, and results with inlet radial distortion showed the expected loss of range. CSTALL was also run in a 3-D mode to investigate inlet circumferential distortion. Calculated operating maps for stage 35 with 120 degree distortion screens showed a loss in range and pressure rise. Unsteady calculations showed rotating stall with two part-span stall cells. The paper describes the body force formulation in detail, examines the computed results, and concludes with observations about the code.

Effect of nanostructures orientation on electroosmotic flow in a microfluidic channel

NASA Astrophysics Data System (ADS)

Eng Lim, An; Lim, Chun Yee; Cheong Lam, Yee; Taboryski, Rafael; Rui Wang, Shu

2017-06-01

Electroosmotic flow (EOF) is an electric-field-induced fluid flow that has numerous micro-/nanofluidic applications, ranging from pumping to chemical and biomedical analyses. Nanoscale networks/structures are often integrated in microchannels for a broad range of applications, such as electrophoretic separation of biomolecules, high reaction efficiency catalytic microreactors, and enhancement of heat transfer and sensing. Their introduction has been known to reduce EOF. Hitherto, a proper study on the effect of nanostructures orientation on EOF in a microfluidic channel is yet to be carried out. In this investigation, we present a novel fabrication method for nanostructure designs that possess maximum orientation difference, i.e. parallel versus perpendicular indented nanolines, to examine the effect of nanostructures orientation on EOF. It consists of four phases: fabrication of silicon master, creation of mold insert via electroplating, injection molding with cyclic olefin copolymer, and thermal bonding and integration of practical inlet/outlet ports. The effect of nanostructures orientation on EOF was studied experimentally by current monitoring method. The experimental results show that nanolines which are perpendicular to the microchannel reduce the EOF velocity significantly (approximately 20%). This flow velocity reduction is due to the distortion of local electric field by the perpendicular nanolines at the nanostructured surface as demonstrated by finite element simulation. In contrast, nanolines which are parallel to the microchannel have no effect on EOF, as it can be deduced that the parallel nanolines do not distort the local electric field. The outcomes of this investigation contribute to the precise control of EOF in lab-on-chip devices, and fundamental understanding of EOF in devices which utilize nanostructured surfaces for chemical and biological analyses.

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.

Hot gas ingestion characteristics and flow visualization of a vectored thrust STOVL concept

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Neiner, George H.; Bencic, Timothy J.; Flood, Joseph D.; Amuedo, Kurt C.; Strock, Thomas W.; Williams, Ben R.

1990-01-01

A 9.2 percent scale short takeoff and vertical landing (STOVL) hot gas ingestion model was designed and built by McDonnell Douglas Corporation (MCAIR) and tested in the NASA Lewis Research Center 9- by 15-Foot Low Speed Wind Tunnel (LSWT). Hot gas ingestion, the entrainment of heated engine exhaust into the inlet flow field, is a key development issue for advanced short takeoff and vertical landing aircraft. The Phase 1 test program, conducted by NASA Lewis and McDonnell Douglas Corporation, evaluated the hot ingestion phenomena and control techniques and Phase 2 test program which was conducted by NASA Lewis are both reported. The Phase 2 program was conducted at exhaust nozzles temperatures up to 1460 R and utilized a sheet laser system for flow visualization of the model flow field in and out of ground effects. Hot gas ingestion levels were measured for the several forward nozzle splay configurations and with flow control/lift improvement devices which reduced the hot gas ingestion. The model support system had four degrees of freedom, heated high pressure air for nozzle flow, and a suction system exhaust for inlet flow. The headwind (freestream) velocity for Phase 1 was varied from 8 to 90 kn, with primary data taken in the 8 to 23 kn headwind velocity range. Phase 2 headwind velocity varied from 10 to 23 kn. Results of both Phase 1 and 2 are presented. A description of the model, facility, a new model support system, and a sheet laser illumination system are also provided. Results are presented over a range of main landing gear height (model height) above the ground plane at a 10 kn headwind velocity. The results contain the compressor face pressure and temperature distortions, total pressure recovery, compressor face temperature rise, and the environmental effects of the hot gas. The environmental effects include the ground plane temperature and pressure distributions, model airframe heating, and the location of the ground flow separation. Results from the sheet laser flow visualization test are also shown.

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

NASA Technical Reports Server (NTRS)

Lawless, Patrick B.; Fleeter, Sanford

1991-01-01

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

Large-Scale Low-Boom Inlet Test Overview

NASA Technical Reports Server (NTRS)

Hirt, Stefanie

2011-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1978-01-01

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

Apparatus and method for continuous production of materials

DOEpatents

Chang, Chih-hung; Jin, Hyungdae

2014-08-12

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

Study of noise sources in a subsonic fan using measured blade pressures and acoustic theory

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1975-01-01

Sources of noise in a 1.4 m (4.6 ft) diameter subsonic tip speed propulsive fan running statically outdoors are studied using a combination of techniques. Signals measured with pressure transducers on a rotor blade are plotted in a format showing the space-time history of inlet distortion. Study of these plots visually and with statistical correlation analysis confirms that the inlet flow contains long, thin eddies of turbulence. Turbulence generated in the boundary layer of the shroud upstream of the rotor tips was not found to be an important noise source. Fan noise is diagnosed by computing narrowband spectra of rotor and stator sound power and comparing these with measured sound power spectra. Rotor noise is computed from spectra of the measured blade pressures and stator noise is computed using the author's stator noise theory. It is concluded that the rotor and stator sources contribute about equally at frequencies in the vicinity of the first three harmonics of blade passing frequency. At higher frequencies, the stator contribution diminishes rapidly and the rotor/inlet turbulence mechanism dominates. Two parametric studies are performed by using the rotor noise calculation procedure which was correlated with test. In the first study, the effects on noise spectrum and directivity are calculated for changes in turbulence properties, rotational Mach number, number of blades, and stagger angle. In the second study the influences of design tip speed and blade number on noise are evaluated.

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

Harmonic engine

DOEpatents

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

2014-08-19

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

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

NASA Technical Reports Server (NTRS)

Beke, Andrew; Allen, J L

1953-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1993-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2014-11-01

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

Optimal Micro-Scale Secondary Flow Control for the Management of High Cycle Fatigue and Distortion in Compact Inlet Diffusers

NASA Technical Reports Server (NTRS)

Anderson, Bernhard H.; Keller, Dennis J.

2002-01-01

The purpose of this study on micro-scale secondary flow control (MSFC) is to study the aerodynamic behavior of micro-vane effectors through their factor (i.e., the design variable) interactions and to demonstrate how these statistical interactions, when brought together in an optimal manner, determine design robustness. The term micro-scale indicates the vane effectors are small in comparison to the local boundary layer height. Robustness in this situation means that it is possible to design fixed MSFC robust installation (i.e.. open loop) which operates well over the range of mission variables and is only marginally different from adaptive (i.e., closed loop) installation design, which would require a control system. The inherent robustness of MSFC micro-vane effector installation designs comes about because of their natural aerodynamic characteristics and the manner in which these characteristics are brought together in an optimal manner through a structured Response Surface Methodology design process.

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.

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.

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.

NASA Lewis Helps Company With New Single-Engine Business Turbojet

NASA Technical Reports Server (NTRS)

1998-01-01

Century Aerospace Corporation, a small company in Albuquerque, New Mexico, is developing a six-seat aircraft powered by a single turbofan engine for general aviation. The company had completed a preliminary design of the jet but needed analyses and testing to proceed with detailed design and subsequent fabrication of a prototype aircraft. NASA Lewis Research Center used computational fluid dynamics (CFD) analyses to ferret out areas of excessive curvature in the inlet where separation might occur. A preliminary look at the results indicated very good inlet performance; and additional calculations, performed with vortex generators installed in the inlet, led to even better results. When it was initially determined that the airflow distortion pattern at the compressor face fell outside of the limits set by the engine manufacturer, the Lewis team studied possible solutions, selected the best, and provided recommendations. CFD results for the inlet system were so good that wind tunnel tests were unnecessary.

Predicting the Inflow Distortion Tone Noise of the NASA Glenn Advanced Noise Control Fan with a Combined Quadrupole-Dipole Model

NASA Technical Reports Server (NTRS)

Koch, L. Danielle

2012-01-01

A combined quadrupole-dipole model of fan inflow distortion tone noise has been extended to calculate tone sound power levels generated by obstructions arranged in circumferentially asymmetric locations upstream of a rotor. Trends in calculated sound power level agreed well with measurements from tests conducted in 2007 in the NASA Glenn Advanced Noise Control Fan. Calculated values of sound power levels radiated upstream were demonstrated to be sensitive to the accuracy of the modeled wakes from the cylindrical rods that were placed upstream of the fan to distort the inflow. Results indicate a continued need to obtain accurate aerodynamic predictions and measurements at the fan inlet plane as engineers work towards developing fan inflow distortion tone noise prediction tools.

Experimental evaluation of a TF30-P-3 turbofan engine in an altitude facility: Effect of steady-state temperature distortion

NASA Technical Reports Server (NTRS)

Braithwaite, W. M.

1973-01-01

The effects of circumferential distortion of the total temperature entering 25, 50, and 75 percent of the inlet circumferential annulus of a turbofan engine were determined. Complete compressor stall resulted from distortions of from 14 to 20 percent of the face averaged temperature. Increasing the temperature level in one sector resulted in that sector moving toward stall by decreasing the equivalent rotor speeds while the pressure ratio remained approximately constant. Stall originated as a rotating zone in the low-pressure compressor which resulted as a terminal stall in the high-pressure compressor. Decreasing the Reynolds number index to 0.25 from 0.5 reduced the required distortion for stall by 50 percent for the conditions investigated.

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

PubMed

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

2014-11-01

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

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

NASA Technical Reports Server (NTRS)

Dennard, John S.

1959-01-01

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

The Effect of Upstream Vane Wakes on Annular Diffuser Flows

NASA Astrophysics Data System (ADS)

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

2008-11-01

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

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

NASA Technical Reports Server (NTRS)

Johns, Albert L.; Burstadt, Paul L.

1984-01-01

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

Fuel cell membrane hydration and fluid metering

DOEpatents

Jones, Daniel O.; Walsh, Michael M.

1999-01-01

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

Navier-Stokes analysis and experimental data comparison of compressible flow within ducts

NASA Technical Reports Server (NTRS)

Harloff, G. J.; Reichert, B. A.; Sirbaugh, J. R.; Wellborn, S. R.

1992-01-01

Many aircraft employ ducts with centerline curvature or changing cross-sectional shape to join the engine with inlet and exhaust components. S-ducts convey air to the engine compressor from the intake and often decelerate the flow to achieve an acceptable Mach number at the engine compressor by increasing the cross-sectional area downstream. Circular-to-rectangular transition ducts are used on aircraft with rectangular exhaust nozzles to connect the engine and nozzle. To achieve maximum engine performance, the ducts should minimize flow total pressure loss and total pressure distortion at the duct exit. Changes in the curvature of the duct centerline or the duct cross-sectional shape give rise to streamline curvature which causes cross stream pressure gradients. Secondary flows can be caused by deflection of the transverse vorticity component of the boundary layer. This vortex tilting results in counter-rotating vortices. Additionally, the adverse streamwise pressure gradient caused by increasing cross-sectional area can lead to flow separation. Vortex pairs have been observed in the exit planes of both duct types. These vortices are due to secondary flows induced by pressure gradients resulting from streamline curvature. Regions of low total pressure are produced when the vortices convect boundary layer fluid into the main flow. The purpose of the present study is to predict the measured flow field in a diffusing S-duct and a circular-to-rectangular transition duct with a full Navier-Stokes computer program, PARC3D, and to compare the numerical predictions with new detailed experimental measurements. The work was undertaken to extend previous studies and to provide additional CFD validation data needed to help model flows with strong secondary flow and boundary layer separation. The S-duct computation extends the study of Smith et al, and Harloff et al, which concluded that the computation might be improved by using a finer grid and more advanced turbulence models. The present study compares results for both the Baldwin-Lomas and k-epsilon turbulence models and is conducted with a refined grid. For the transition duct, two inlet conditions were considered, the first with straight flow and the second with swirling flow. The first case permits examination of the effects of the geometric transition on the flow field, while the second case includes the rotational flow effect characteristic of a gas turbine engine.

Radial lean direct injection burner

DOEpatents

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

2012-09-04

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

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

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

Boosting devices with integral features for recirculating exhaust gas

DOEpatents

Wu, Ko -Jen

2015-09-15

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

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

NASA Astrophysics Data System (ADS)

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

2013-12-01

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

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

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

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

1976-01-01

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

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

NASA Technical Reports Server (NTRS)

Porro, A. Robert

2000-01-01

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

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)

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.

Rotary-Wing Relevant Compressor Aero Research and Technology Development Activities at Glenn Research Center

NASA Technical Reports Server (NTRS)

Welch, Gerard E.; Hathaway, Michael D.; Skoch, Gary J.; Snyder, Christopher A.

2012-01-01

Technical challenges of compressors for future rotorcraft engines are driven by engine-level and component-level requirements. Cycle analyses are used to highlight the engine-level challenges for 3000, 7500, and 12000 SHP-class engines, which include retention of performance and stability margin at low corrected flows, and matching compressor type, axial-flow or centrifugal, to the low corrected flows and high temperatures in the aft stages. At the component level: power-to-weight and efficiency requirements impel designs with lower inherent aerodynamic stability margin; and, optimum engine overall pressure ratios lead to small blade heights and the associated challenges of scale, particularly increased clearance-to-span ratios. The technical challenges associated with the aerodynamics of low corrected flows and stability management impel the compressor aero research and development efforts reviewed herein. These activities include development of simple models for clearance sensitivities to improve cycle calculations, full-annulus, unsteady Navier-Stokes simulations used to elucidate stall, its inception, and the physics of stall control by discrete tip-injection, development of an actuator-duct-based model for rapid simulation of nonaxisymmetric flow fields (e.g., due inlet circumferential distortion), advanced centrifugal compressor stage development and experimentation, and application of stall control in a T700 engine.

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

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2011-01-01

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

Investigation of the effects of inlet shapes on fan noise radiation

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

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.

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.

Quasi One-Dimensional Unsteady Modeling of External Compression Supersonic Inlets

NASA Technical Reports Server (NTRS)

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

2012-01-01

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

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

NASA Technical Reports Server (NTRS)

Agnone, Anthony M.

1987-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

Miniature piezo electric vacuum inlet valve

DOEpatents

Keville, Robert F.; Dietrich, Daniel D.

1998-03-24

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

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

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

Aerodynamic design of gas and aerosol samplers for aircraft

NASA Technical Reports Server (NTRS)

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

1991-01-01

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

SUPIN: A Computational Tool for Supersonic Inlet Design

NASA Technical Reports Server (NTRS)

Slater, John W.

2016-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1994-01-01

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

Systems to facilitate reducing flashback/flame holding in combustion systems

DOEpatents

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

2012-02-21

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

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

NASA Technical Reports Server (NTRS)

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

1988-01-01

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

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

NASA Astrophysics Data System (ADS)

Gao, WenZhi; Li, ZhuFei; Yang, JiMing

2015-10-01

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

The performance of a centrifugal compressor with high inlet prewhirl

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

Whitfield, A.; Abdullah, A.H.

1998-07-01

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

Adaptive compliant structures for flow regulation

PubMed Central

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

2017-01-01

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

Adaptive compliant structures for flow regulation.

PubMed

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

2017-08-01

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

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

NASA Technical Reports Server (NTRS)

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

2009-01-01

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

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

NASA Astrophysics Data System (ADS)

Du, B. Y.

2016-10-01

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

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

NASA Technical Reports Server (NTRS)

Hearth, Donald P; Cubbison, Robert W

1956-01-01

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

Removable feedwater sparger assembly

DOEpatents

Challberg, R.C.

1994-10-04

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

Removable feedwater sparger assembly

DOEpatents

Challberg, Roy C.

1994-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1947-01-01

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

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

NASA Astrophysics Data System (ADS)

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

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

Validation of WIND for a Series of Inlet Flows

NASA Technical Reports Server (NTRS)

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

2002-01-01

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

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

NASA Astrophysics Data System (ADS)

Zhang, Zh.

2018-02-01

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

Entropy considerations applied to shock unsteadiness in hypersonic inlets

NASA Astrophysics Data System (ADS)

Bussey, Gillian Mary Harding

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

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

NASA Astrophysics Data System (ADS)

Nilsson, H.; Cervantes, M. J.

2012-11-01

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

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

NASA Astrophysics Data System (ADS)

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

2013-10-01

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

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

NASA Technical Reports Server (NTRS)

Re, R. J.

1974-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1974-01-01

The flow field of a jet engine with an inlet vortex, which can pick up heavy debris from the ground and damage the engine, was simulated in a small water tunnel by means of the hydrogen bubble technique. It was found that the known engine inlet vortex is accompained by a vortex system, consisting of two inlet vortices (the ground based and the trailing one), secondary vortices, and ground vortices. Simulation of the ground effect by an inlet image proved that the inlet vortex feeds on free stream vorticity and can exist without the presence of a ground boundary layer. The structural form of the inlet vortex system was explained by a simple potential flow model, which showed the number, location, and the importance of the stagnation points. A retractable horizontal screen or an up-tilt of the engine is suggested as countermeasure against debris ingestion.

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.

An Optical Flow-Based Full Reference Video Quality Assessment Algorithm.

PubMed

K, Manasa; Channappayya, Sumohana S

2016-06-01

We present a simple yet effective optical flow-based full-reference video quality assessment (FR-VQA) algorithm for assessing the perceptual quality of natural videos. Our algorithm is based on the premise that local optical flow statistics are affected by distortions and the deviation from pristine flow statistics is proportional to the amount of distortion. We characterize the local flow statistics using the mean, the standard deviation, the coefficient of variation (CV), and the minimum eigenvalue ( λ min ) of the local flow patches. Temporal distortion is estimated as the change in the CV of the distorted flow with respect to the reference flow, and the correlation between λ min of the reference and of the distorted patches. We rely on the robust multi-scale structural similarity index for spatial quality estimation. The computed temporal and spatial distortions, thus, are then pooled using a perceptually motivated heuristic to generate a spatio-temporal quality score. The proposed method is shown to be competitive with the state-of-the-art when evaluated on the LIVE SD database, the EPFL Polimi SD database, and the LIVE Mobile HD database. The distortions considered in these databases include those due to compression, packet-loss, wireless channel errors, and rate-adaptation. Our algorithm is flexible enough to allow for any robust FR spatial distortion metric for spatial distortion estimation. In addition, the proposed method is not only parameter-free but also independent of the choice of the optical flow algorithm. Finally, we show that the replacement of the optical flow vectors in our proposed method with the much coarser block motion vectors also results in an acceptable FR-VQA algorithm. Our algorithm is called the flow similarity index.

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.

Miniature piezo electric vacuum inlet valve

DOEpatents

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

1998-03-24

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

Large perturbation flow field analysis and simulation for supersonic inlets

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

Computer programs for calculating potential flow in propulsion system inlets

NASA Technical Reports Server (NTRS)

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

1973-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1983-01-01

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

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

NASA Technical Reports Server (NTRS)

Brown, David R.; Brown, Kenneth G.

1988-01-01

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

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.

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.

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

PubMed

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

2015-06-01

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

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

DOEpatents

Kozubal, Eric Joseph

2016-12-13

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

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

DOEpatents

Kozubal, Eric Joseph; Slayzak, Steven Joseph

2014-07-08

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

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.

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

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

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

Three Dimensional Compressible Turbulent Flow Computations for a Diffusing S-Duct With/Without Vortex Generators

NASA Technical Reports Server (NTRS)

Cho, Soo-Yong; Greber, Isaac

1994-01-01

Numerical investigations on a diffusing S-duct with/without vortex generators and a straight duct with vortex generators are presented. The investigation consists of solving the full three-dimensional unsteady compressible mass averaged Navier-Stokes equations. An implicit finite volume lower-upper time marching code (RPLUS3D) has been employed and modified. A three-dimensional Baldwin-Lomax turbulence model has been modified in conjunction with the flow physics. A model for the analysis of vortex generators in a fully viscous subsonic internal flow is evaluated. A vortical structure for modeling the shed vortex is used as a source term in the computation domain. The injected vortex paths in the straight duct are compared with the analysis by two kinds of prediction models. The flow structure by the vortex generators are investigated along the duct. Computed results of the flow in a circular diffusing S-duct provide an understanding of the flow structure within a typical engine inlet system. These are compared with the experimental wall static-pressure, static- and total-pressure field, and secondary velocity profiles. Additionally, boundary layer thickness, skin friction values, and velocity profiles in wall coordinates are presented. In order to investigate the effect of vortex generators, various vortex strengths are examined in this study. The total-pressure recovery and distortion coefficients are obtained at the exit of the S-duct. The numerical results clearly depict the interaction between the low velocity flow by the flow separation and the injected vortices.

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

DTIC Science & Technology

2010-01-01

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

Polyport atmospheric gas sampler

DOEpatents

Guggenheim, S. Frederic

1995-01-01

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

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

NASA Technical Reports Server (NTRS)

Benson, Thomas J.

1988-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2009-01-01

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

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

PubMed

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

2016-12-12

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

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

NASA Technical Reports Server (NTRS)

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

1990-01-01

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

Inlet Flow Valve Engine Analyses

NASA Technical Reports Server (NTRS)

Champagne, G. A.

2004-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

Forebody and Inlet Design for the HIFiRE 2 Flight Test

NASA Technical Reports Server (NTRS)

Ferlemann, Paul G.

2008-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1981-01-01

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

Flow Control in a Compact Inlet

NASA Astrophysics Data System (ADS)

Vaccaro, John C.

2011-12-01

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

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

NASA Technical Reports Server (NTRS)

Pendley, Robert E; Robinson, Harold L

1950-01-01

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

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

NASA Astrophysics Data System (ADS)

Barmina, I.; Valdmanis, R.; Zaķe, M.

2017-06-01

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

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

NASA Technical Reports Server (NTRS)

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

1999-01-01

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

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.

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

USGS Publications Warehouse

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

2013-01-01

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

Combined Cycle Engine Large-Scale Inlet for Mode Transition Experiments: System Identification Rack Hardware Design

NASA Technical Reports Server (NTRS)

Thomas, Randy; Stueber, Thomas J.

2013-01-01

The System Identification (SysID) Rack is a real-time hardware-in-the-loop data acquisition (DAQ) and control instrument rack that was designed and built to support inlet testing in the NASA Glenn Research Center 10- by 10-Foot Supersonic Wind Tunnel. This instrument rack is used to support experiments on the Combined-Cycle Engine Large-Scale Inlet for Mode Transition Experiment (CCE? LIMX). The CCE?LIMX is a testbed for an integrated dual flow-path inlet configuration with the two flow paths in an over-and-under arrangement such that the high-speed flow path is located below the lowspeed flow path. The CCE?LIMX includes multiple actuators that are designed to redirect airflow from one flow path to the other; this action is referred to as "inlet mode transition." Multiple phases of experiments have been planned to support research that investigates inlet mode transition: inlet characterization (Phase-1) and system identification (Phase-2). The SysID Rack hardware design met the following requirements to support Phase-1 and Phase-2 experiments: safely and effectively move multiple actuators individually or synchronously; sample and save effector control and position sensor feedback signals; automate control of actuator positioning based on a mode transition schedule; sample and save pressure sensor signals; and perform DAQ and control processes operating at 2.5 KHz. This document describes the hardware components used to build the SysID Rack including their function, specifications, and system interface. Furthermore, provided in this document are a SysID Rack effectors signal list (signal flow); system identification experiment setup; illustrations indicating a typical SysID Rack experiment; and a SysID Rack performance overview for Phase-1 and Phase-2 experiments. The SysID Rack described in this document was a useful tool to meet the project objectives.

Experimental Evaluation of Turning Vane Designs for High-speed and Coupled Fan-drive Corners of 0.1-scale Model of NASA Lewis Research Center's Proposed Altitude Wind Tunnel

NASA Technical Reports Server (NTRS)

Gelder, Thomas F.; Moore, Royce D.; Shyne, Rickey J.; Boldman, Donald R.

1987-01-01

Two turning vane designs were experimentally evaluated for the fan-drive corner (corner 2) coupled to an upstream diffuser and the high-speed corner (corner 1) of the 0.1 scale model of NASA Lewis Research Center's proposed Altitude Wind Tunnel. For corner 2 both a controlled-diffusion vane design (vane A4) and a circular-arc vane design (vane B) were studied. The corner 2 total pressure loss coefficient was about 0.12 with either vane design. This was about 25 percent less loss than when corner 2 was tested alone. Although the vane A4 design has the advantage of 20 percent fewer vanes than the vane B design, its vane shape is more complex. The effects of simulated inlet flow distortion on the overall losses for corner 1 or 2 were small.

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

NASA Astrophysics Data System (ADS)

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

2015-08-01

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

IPAC-Inlet Performance Analysis Code

NASA Technical Reports Server (NTRS)

Barnhart, Paul J.

1997-01-01

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

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

NASA Technical Reports Server (NTRS)

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

2011-01-01

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

Study of noise and inflow distortion sources in the NASA QF-1B fan using measured blade and vane pressures

NASA Technical Reports Server (NTRS)

Hanson, D. B.

1977-01-01

Pressure transducers were installed on the blades and vanes of QF-1B, a transonic-tip-speed fan from the NASA Quiet Fan Program which was tested on the outdoor quiet fan test facility at NASA-Lewis. Signals from the transducers and from far field microphones were analyzed to determine sources of nonuniform inflow and noise. The nonuniform inflow was mostly unsteady with roughly equal contributions from atmospheric turbulence and rig interference. The rig interference was largest at the bottom and appeared to be generated by the support structure which was located behind the inlet lip under the fan. Interaction of this inflow distortion was the dominant source of noise at 1, 2, and 3 times blade passing frequency (BPF) at 60, 70, and 80 percent of design speed. At 90 percent speed, noise at BPF was dominated by the steady rotor field. A broadband spectrum peak centered at about 2.2 times BPF was identified as rotor/stator interaction stemming from a high frequency rotor exit flow component. The remaining broadband energy from 0.3 to 3.5 times BPF was attributed to the better known type of rotor/stator interaction associated with rotor wake turbulence.

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

NASA Astrophysics Data System (ADS)

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

2018-04-01

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

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.

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

Acoustically enhanced heat exchange and drying apparatus

DOEpatents

Bramlette, T. Tazwell; Keller, Jay O.

1989-01-01

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

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

NASA Technical Reports Server (NTRS)

Stabe, Roy G.; Schwab, John R.

1991-01-01

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

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

NASA Technical Reports Server (NTRS)

Torrence, M. G.

1975-01-01

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

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

NASA Technical Reports Server (NTRS)

Albers, J. A.

1973-01-01

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

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

PubMed Central

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

2011-01-01

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

Supersonic Elliptical Ramp Inlet

NASA Technical Reports Server (NTRS)

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

2016-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

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

NASA Technical Reports Server (NTRS)

Tornabene, Robert

2005-01-01

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

Electric fluid pump

DOEpatents

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

2015-09-29

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

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

NASA Technical Reports Server (NTRS)

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

1976-01-01

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

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

NASA Technical Reports Server (NTRS)

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

1984-01-01

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

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

DOEpatents

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

2009-03-03

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

Detailed design of a quiet high flow fan

NASA Technical Reports Server (NTRS)

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

1977-01-01

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

Unstart phenomena induced by flow choking in scramjet inlet-isolators

NASA Astrophysics Data System (ADS)

Im, Seong-kyun; Do, Hyungrok

2018-02-01

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

Flow behavior in inlet guide vanes of radial turbines

NASA Technical Reports Server (NTRS)

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

1975-01-01

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

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

DTIC Science & Technology

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

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

NASA Technical Reports Server (NTRS)

Lewis, James P.

1953-01-01

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

Development of Wing Inlets

NASA Technical Reports Server (NTRS)

Racisz, Stanley F.

1946-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2018-05-01

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

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

PubMed Central

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

2014-01-01

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

Method for Determining Optimum Injector Inlet Geometry

NASA Technical Reports Server (NTRS)

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

2015-01-01

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

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

NASA Astrophysics Data System (ADS)

Abdel-Hameed, H.; Bellan, J.

2002-10-01

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

Analytical models for use in fan inflow control structure design. Inflow distortion and acoustic transmission models

NASA Technical Reports Server (NTRS)

Gedge, M. R.

1979-01-01

Analytical models were developed to study the effect of flow contraction and screening on inflow distortions to identify qualitative design criteria. Results of the study are that: (1) static testing distortions are due to atmospheric turbulence, nacelle boundary layer, exhaust flow reingestion, flow over stand, ground plane, and engine casing; (2) flow contraction suppresses, initially, turbulent axial velocity distortions and magnifies turbulent transverse velocity distortions; (3) perforated plate and gauze screens suppress axial components of velocity distortions to a degree determined by the screen pressure loss coefficient; (4) honeycomb screen suppress transverse components of velocity distortions to a degree determined by the length to diameter ratio of the honeycomb; (5) acoustic transmission loss of perforated plate is controlled by the reactance of its acoustic impedance; (6) acoustic transmission loss of honeycomb screens is negligible; and (7) a model for the direction change due to a corner between honeycomb panels compares favorably with measured data.

Axial Flow Conditioning Device for Mitigating Instabilities

NASA Technical Reports Server (NTRS)

Ahuja, Vineet (Inventor); Birkbeck, Roger M. (Inventor); Hosangadi, Ashvin (Inventor)

2017-01-01

A flow conditioning device for incrementally stepping down pressure within a piping system is presented. The invention includes an outer annular housing, a center element, and at least one intermediate annular element. The outer annular housing includes an inlet end attachable to an inlet pipe and an outlet end attachable to an outlet pipe. The outer annular housing and the intermediate annular element(s) are concentrically disposed about the center element. The intermediate annular element(s) separates an axial flow within the outer annular housing into at least two axial flow paths. Each axial flow path includes at least two annular extensions that alternately and locally direct the axial flow radially outward and inward or radially inward and outward thereby inducing a pressure loss or a pressure gradient within the axial flow. The pressure within the axial flow paths is lower than the pressure at the inlet end and greater than the vapor pressure for the axial flow. The invention minimizes fluidic instabilities, pressure pulses, vortex formation and shedding, and/or cavitation during pressure step down to yield a stabilized flow within a piping system.

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

NASA Technical Reports Server (NTRS)

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

1985-01-01

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

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

NASA Astrophysics Data System (ADS)

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

2012-07-01

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

Calculation of compressible flow in and about three-dimensional inlets with and without auxiliary inlets by a higher-order panel method

NASA Technical Reports Server (NTRS)

Hess, J. L.; Friedman, D. M.

1982-01-01

A three dimensional higher order panel method was specialized to the case of inlets with auxiliary inlets. The resulting program has a number of graphical input-output features to make it highly useful to the designer. The various aspects of the program are described instructions for its use are presented.

An experimental and computational investigation of flow in a radial inlet of an industrial pipeline centrifugal compressor

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

Flathers, M.B.; Bache, G.E.; Rainsberger, R.

1996-04-01

The flow field of a complex three-dimensional radial inlet for an industrial pipeline centrifugal compressor has been experimentally determined on a half-scale model. Based on the experimental results, inlet guide vanes have been designed to correct pressure and swirl angle distribution deficiencies. The unvaned and vaned inlets are analyzed with a commercially available fully three-dimensional viscous Navier-Stokes code. Since experimental results were available prior to the numerical study, the unvaned analysis is considered a postdiction while the vaned analysis is considered a prediction. The computational results of the unvaned inlet have been compared to the previously obtained experimental results. Themore » experimental method utilized for the unvaned inlet is repeated for the vaned inlet and the data have been used to verify the computational results. The paper will discuss experimental, design, and computational procedures, grid generation, boundary conditions, and experimental versus computational methods. Agreement between experimental and computational results is very good, both in prediction and postdiction modes. The results of this investigation indicate that CFD offers a measurable advantage in design, schedule, and cost and can be applied to complex, three-dimensional radial inlets.« less

77 FR 30283 - Standards of Performance for New Stationary Sources, National Emission Standards for Hazardous...

Federal Register 2010, 2011, 2012, 2013, 2014

2012-05-22

... timeline and exceedances of the pressure, temperature, and oxygen and/or nitrogen concentration are... (AMP) that consists of monitoring the inlet scrubbing liquid temperature, flow rate, and acid content... request consisting of monitoring the inlet scrubbing liquid temperature and flow rate and identifying...

Remote actuated valve implant

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

McKnight, Timothy E.; Johnson, Anthony; Moise, Kenneth J.

Valve implant systems positionable within a flow passage, the systems having an inlet, an outlet, and a remotely activatable valve between the inlet and outlet, with the valves being operable to provide intermittent occlusion of the flow path. A remote field is applied to provide thermal or magnetic activation of the valves.

Analytical and Experimental Study of Flow Through an Axial Turbine Stage with a Nonuniform Inlet Radial Temperature Profile

NASA Technical Reports Server (NTRS)

Schwab, J. R.; Stabe, R. G.; Whitney, W. J.

1983-01-01

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects.

Analytical and experimental study of flow through an axial turbine stage with a nonuniform inlet radial temperature profile

NASA Technical Reports Server (NTRS)

Schwab, J. R.; Stabe, R. G.; Whitney, W. J.

1983-01-01

Results are presented for a typical nonuniform inlet radial temperature profile through an advanced single-stage axial turbine and compared with the results obtained for a uniform profile. Gas temperature rises of 40 K to 95 K are predicted at the hub and tip corners at the trailing edges of the pressure surfaces in both the stator and rotor due to convection of hot fluid from the mean by the secondary flow. The inlet temperature profile is shown to be mixed out at the rotor exit survey plane (2.3 axial chords downstream of the rotor trailing edge) in both the analysis and the experiment. The experimental rotor exit angle profile for the nonuniform inlet temperature profile indicates underturning at the tip caused by increased clearance. Severe underturning also occurs at the mean, both with and without the nonuniform inlet temperature profile. The inviscid rotational flow code used in the analysis fails to predict the underturning at the mean, which may be caused by viscous effects. Previously announced in STAR as N83-27958

Preliminary Investigation of a New Type of Supersonic Inlet

NASA Technical Reports Server (NTRS)

Ferri, Antonio; Nucci, Louis M

1946-01-01

A supersonic inlet with supersonic deceleration of the flow entirely outside of the inlet is considered. A particular arrangement with fixed geometry having a central body with a circular annular intake is analyzed, and it is shown theoretically that this arrangement gives high pressure recovery for a large range of Mach number and mass flow and therefore is practical for use on supersonic airplanes and missiles. For some Mach numbers the drag coefficient for this type of inlet is larger than the drag coefficient for the type of inlet with supersonic compression entirely inside, but the pressure recovery is larger for all flight conditions. The differences in drag can be eliminated for the design Mach number. Experimental results confirm the results of the theoretical analysis and show that pressure recoveries of 95 percent for Mach numbers of 1.33 and 1.52, 92 percent for a Mach number of 1.72, and 86 percent for a Mach number oof 2.10 are possible with the configurations considered. If the mass flow decreases, the total drag coefficient increases gradually and the pressure recovery does not change appreciably.

Supersonic compressor

DOEpatents

Lawlor, Shawn P [Bellevue, WA; Novaresi, Mark A [San Diego, CA; Cornelius, Charles C [Kirkland, WA

2008-02-26

A gas compressor based on the use of a driven rotor having an axially oriented compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which forms a supersonic shockwave axially, between adjacent strakes. In using this method to compress inlet gas, the supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdyanamic flow path formed between the gas compression ramp on a strake, the shock capture lip on the adjacent strake, and captures the resultant pressure within the stationary external housing while providing a diffuser downstream of the compression ramp.

Solar Cell Modules With Improved Backskin

DOEpatents

Chevrefils, Andre; Grigore, Daniel Gheorghe

2001-01-23

The present invention relates to gas turbines and more particularly to a device for controlling the flow of cooling air through a flowpath in a turbine blade. The device can be inserted in the inlet opening of the blade flowpath and be retained therein. The device comprises a plug member for adjusting the flow of cooling air through the flowpath. The plug member comprises a retaining portion for retaining the plug member at the inlet opening of the flowpath and a blocking portion inserted within the flowpath for reducing the cross-sectional area of the inlet opening. Such a device is inexpensive and can be easily inserted in the inlet opening of a blade flowpath and retained therein.

A Framework for Simulating Turbine-Based Combined-Cycle Inlet Mode-Transition

NASA Technical Reports Server (NTRS)

Le, Dzu K.; Vrnak, Daniel R.; Slater, John W.; Hessel, Emil O.

2012-01-01

A simulation framework based on the Memory-Mapped-Files technique was created to operate multiple numerical processes in locked time-steps and send I/O data synchronously across to one-another to simulate system-dynamics. This simulation scheme is currently used to study the complex interactions between inlet flow-dynamics, variable-geometry actuation mechanisms, and flow-controls in the transition from the supersonic to hypersonic conditions and vice-versa. A study of Mode-Transition Control for a high-speed inlet wind-tunnel model with this MMF-based framework is presented to illustrate this scheme and demonstrate its usefulness in simulating supersonic and hypersonic inlet dynamics and controls or other types of complex systems.

Fluidic Oscillator Having Decoupled Frequency and Amplitude Control

NASA Technical Reports Server (NTRS)

Koklu, Mehti (Inventor)

2017-01-01

A fluidic oscillator having independent frequency and amplitude control includes a fluidic-oscillator main flow channel having a main flow inlet, a main flow outlet, and first and second control ports disposed at opposing sides thereof. A fluidic-oscillator controller has an inlet and outlet. A volume defined by the main flow channel is greater than the volume defined by the controller. A flow diverter coupled to the outlet of the controller defines a first fluid flow path from the controller's outlet to the first control port and defines a second fluid flow path from the controller's outlet to the second control port.

Fluidic Oscillator Having Decoupled Frequency and Amplitude Control

NASA Technical Reports Server (NTRS)

Koklu, Mehti (Inventor)

2016-01-01

A fluidic oscillator having independent frequency and amplitude control includes a fluidic-oscillator main flow channel having a main flow inlet, a main flow outlet, and first and second control ports disposed at opposing sides thereof. A fluidic-oscillator controller has an inlet and outlet. A volume defined by the main flow channel is greater than the volume defined by the controller. A flow diverter coupled to the outlet of the controller defines a first fluid flow path from the controller's outlet to the first control port and defines a second fluid flow path from the controller's outlet to the second control port.

The effect of circumferential distortion on fan performance at two levels of blade loading

NASA Technical Reports Server (NTRS)

Hartmann, M. J.; Sanger, N. L.

1975-01-01

Single stage fans designed for two levels of pressure ratio or blade loading were subjected to screen-induced circumferential distortions of 90-degree extent. Both fan rotors were designed for a blade tip speed of 425 m/sec, blade solidity of 1.3 and a hub-to-tip radius ratio of 0.5. Circumferential measurements of total pressure, temperature, static pressure, and flow angle were obtained at the hub, mean and tip radii at five axial stations. Rotor loading level did not appear to have a significant influence on rotor response to distorted flow. Losses in overall pressure ratio due to distortion were most severe in the stator hub region of the more highly loaded stage. At the near stall operating condition tip and hub regions of (either) rotor demonstrated different response characteristics to the distorted flow. No effect of loading was apparent on interactions between rotor and upstream distorted flow fields.

DESIGN ANALYSIS OF RADIAL INFLOW TURBINES

NASA Technical Reports Server (NTRS)

Glassman, A. J.

1994-01-01

This program performs a velocity-diagram analysis required for determining geometry and estimating performance for radial-inflow turbines. Input design requirements are power, mass flow rate, inlet temperature and pressure, and rotative rate. The design variables include stator-exit angle, rotor-exit-tip to rotor-inlet radius ratio, rotor-exit-hub to tip radius ratio, and the magnitude and radial distribution of rotor-exit tangential velocity. The program output includes diameters, total and static efficiences, all absolute and relative temperatures, pressures, and velocities, and flow angles at stator inlet, stator exit, rotor inlet, and rotor exit. Losses accounted for in this program by the internal loss model are three-dimensional (profile plus end wall) viscous losses in the stator and the rotor, the disk-friction loss on the back side of the rotor, the loss due to the clearance between the rotor tip and the outer casing, and the exit velocity loss. The flow analysis is one-dimensional at the stator inlet, stator exit, and rotor inlet, each of these calculation stations being at a constant radius. At the rotor exit where there is a variation in flow-field radius, an axisymmetric two-dimensional analysis is made using constant height sectors. Simple radial equilibrium is used to establish the static pressure gradient at the rotor exit. This program is written in FORTRAN V and has been implemented on a UNIVAC 1100 series computer with a memory requirement of approximately 22K of 36 bit words.

Inward-Turning Streamline-Traced Inlet Design Method for Low-Boom, Low-Drag Applications

NASA Technical Reports Server (NTRS)

Otto, Samuel; Trefny, Charles J.; Slater, John W.

2015-01-01

A new design method for inward-turning, streamline-traced inlets is presented. Resulting designs are intended for moderate supersonic, low-drag, low-boom applications such as that required for NASA's proposed low-boom flight demonstration aircraft. A critical feature of these designs is the internal cowl lip angle that allows for little or no flow turning on the outer nacelle. Present methods using conical-flow Busemann parent flowfields have simply truncated, or otherwise modified the stream-traced contours to include this internal cowl angle. Such modifications disrupt the parent flowfield, reducing inlet performance and flow uniformity. The method presented herein merges a conical flowfield that includes a leading shock with a truncated Busemann flowfield in a manner that minimizes unwanted interactions. A leading internal cowl angle is now inherent in the parent flowfield, and inlet contours traced from this flowfield retain its high performance and good flow uniformity. CFD analysis of a candidate inlet design is presented that verifies the design technique, and reveals a starting issue with the basic geometry. A minor modification to the cowl lip region is shown to eliminate this phenomenon, thereby allowing starting and smooth transition to sub-critical operation as back-pressure is increased. An inlet critical-point total pressure recovery of 96 is achieved based on CFD results for a Mach 1.7 freestream design. Correction for boundary-layer displacement thickness, and sizing for a given engine airflow requirement are also discussed.

Confined turbulent swirling recirculating flow predictions. Ph.D. Thesis. Final Report

NASA Technical Reports Server (NTRS)

Abujelala, M. T.; Lilley, D. G.

1985-01-01

The capability and the accuracy of the STARPIC computer code in predicting confined turbulent swirling recirculating flows is presented. Inlet flow boundary conditions were demonstrated to be extremely important in simulating a flowfield via numerical calculations. The degree of swirl strength and expansion ratio have strong effects on the characteristics of swirling flow. In a nonswirling flow, a large corner recirculation zone exists in the flowfield with an expansion ratio greater than one. However, as the degree of inlet swirl increases, the size of this zone decreases and a central recirculation zone appears near the inlet. Generally, the size of the central zone increased with swirl strength and expansion ratio. Neither the standard k-epsilon turbulence mode nor its previous extensions show effective capability for predicting confined turbulent swirling recirculating flows. However, either reduced optimum values of three parameters in the mode or the empirical C sub mu formulation obtained via careful analysis of available turbulence measurements, can provide more acceptable accuracy in the prediction of these swirling flows.

Aerosol mobility size spectrometer

DOEpatents

Wang, Jian; Kulkarni, Pramod

2007-11-20

A device for measuring aerosol size distribution within a sample containing aerosol particles. The device generally includes a spectrometer housing defining an interior chamber and a camera for recording aerosol size streams exiting the chamber. The housing includes an inlet for introducing a flow medium into the chamber in a flow direction, an aerosol injection port adjacent the inlet for introducing a charged aerosol sample into the chamber, a separation section for applying an electric field to the aerosol sample across the flow direction and an outlet opposite the inlet. In the separation section, the aerosol sample becomes entrained in the flow medium and the aerosol particles within the aerosol sample are separated by size into a plurality of aerosol flow streams under the influence of the electric field. The camera is disposed adjacent the housing outlet for optically detecting a relative position of at least one aerosol flow stream exiting the outlet and for optically detecting the number of aerosol particles within the at least one aerosol flow stream.

Flow control in axial fan inlet guide vanes by synthetic jets

NASA Astrophysics Data System (ADS)

Cyrus, V.; Trávníček, Z.; Wurst, P.; Kordík, J.

2013-04-01

Tested high pressure axial flow fan with hub/tip ratio of 0.70 and external diameter of 600 mm consisted of inlet guide vanes (IGV), rotor and stator blade rows. Fan peripheral velocity was 47 m/s. Air volume flow rate was changed by turning of rear part of the inlet guide vanes. At turning of 20 deg the flow was separated on the IGV profiles. The synthetic jets were introduced through radial holes in machine casing in the location before flow separation origin. Synthetic jet actuator was designed with the use of a speaker by UT AVCR. Its membrane had diameter of 63 mm. Excitation frequency was chosen in the range of 500 Hz - 700 Hz. Synthetic jets favourably influenced separated flow on the vane profiles in the distance of (5 - 12) mm from the casing surface. The reduction of flow separation area caused in the region near the casing the decrease of the profile loss coefficient approximately by 20%.

Design and Analysis Tools for Supersonic Inlets

NASA Technical Reports Server (NTRS)

Slater, John W.; Folk, Thomas C.

2009-01-01

Computational tools are being developed for the design and analysis of supersonic inlets. The objective is to update existing tools and provide design and low-order aerodynamic analysis capability for advanced inlet concepts. The Inlet Tools effort includes aspects of creating an electronic database of inlet design information, a document describing inlet design and analysis methods, a geometry model for describing the shape of inlets, and computer tools that implement the geometry model and methods. The geometry model has a set of basic inlet shapes that include pitot, two-dimensional, axisymmetric, and stream-traced inlet shapes. The inlet model divides the inlet flow field into parts that facilitate the design and analysis methods. The inlet geometry model constructs the inlet surfaces through the generation and transformation of planar entities based on key inlet design factors. Future efforts will focus on developing the inlet geometry model, the inlet design and analysis methods, a Fortran 95 code to implement the model and methods. Other computational platforms, such as Java, will also be explored.

YF-12 Experiments Symposium, Volume 1

NASA Technical Reports Server (NTRS)

1978-01-01

Papers presented by personnel from the Dryden Flight Research Center, the Lewis Research Center, and the Ames Research Center are presented. Topics cover propulsion system performance, inlet time varying distortion, structures, aircraft controls, propulsion controls, and aerodynamics. The reports were based on analytical studies, laboratory experiments, wind tunnel tests, and extensive flight research with two YF-12 airplanes.

Investigation of Blade Angle of an Open Cross-Flow Runner

NASA Astrophysics Data System (ADS)

Katayama, Yusuke; Iio, Shouichiro; Veerapun, Salisa; Uchiyama, Tomomi

2015-04-01

The aim of this study was to develop a nano-hydraulic turbine utilizing drop structure in irrigation channels or industrial waterways. This study was focused on an open-type cross-flow turbine without any attached equipment for cost reduction and easy maintenance. In this study, the authors used an artificial indoor waterfall as lab model. Test runner which is a simple structure of 20 circular arc-shaped blades sandwiched by two circular plates was used The optimum inlet blade angle and the relationship between the power performance and the flow rate approaching theoretically and experimentally were investigated. As a result, the optimum inlet blade angle due to the flow rate was changed. Additionally, allocation rate of power output in 1st stage and 2nd stage is changed by the blade inlet angle.

Experimental study of the influence of flow passage subtle variation on mixed-flow pump performance

NASA Astrophysics Data System (ADS)

Bing, Hao; Cao, Shuliang

2014-05-01

In the mixed-flow pump design, the shape of the flow passage can directly affect the flow capacity and the internal flow, thus influencing hydraulic performance, cavitation performance and operation stability of the mixed-flow pump. However, there is currently a lack of experimental research on the influence mechanism. Therefore, in order to analyze the effects of subtle variations of the flow passage on the mixed-flow pump performance, the frustum cone surface of the end part of inlet contraction flow passage of the mixed-flow pump is processed into a cylindrical surface and a test rig is built to carry out the hydraulic performance experiment. In this experiment, parameters, such as the head, the efficiency, and the shaft power, are measured, and the pressure fluctuation and the noise signal are also collected. The research results suggest that after processing the inlet flow passage, the head of the mixed-flow pump significantly goes down; the best efficiency of the mixed-flow pump drops by approximately 1.5%, the efficiency decreases more significantly under the large flow rate; the shaft power slightly increases under the large flow rate, slightly decreases under the small flow rate. In addition, the pressure fluctuation amplitudes on both the impeller inlet and the diffuser outlet increase significantly with more drastic pressure fluctuations and significantly lower stability of the internal flow of the mixed-flow pump. At the same time, the noise dramatically increases. Overall speaking, the subtle variation of the inlet flow passage leads to a significant change of the mixed-flow pump performance, thus suggesting a special attention to the optimization of flow passage. This paper investigates the influence of the flow passage variation on the mixed-flow pump performance by experiment, which will benefit the optimal design of the flow passage of the mixed-flow pump.

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

NASA Technical Reports Server (NTRS)

Huebner, Lawrence D.; Tatum, Kenneth E.

1993-01-01

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

Comparison Between Numerically Simulated and Experimentally Measured Flowfield Quantities Behind a Pulsejet

NASA Technical Reports Server (NTRS)

Geng, Tao; Paxson, Daniel E.; Zheng, Fei; Kuznetsov, Andrey V.; Roberts, William L.

2008-01-01

Pulsed combustion is receiving renewed interest as a potential route to higher performance in air breathing propulsion systems. Pulsejets offer a simple experimental device with which to study unsteady combustion phenomena and validate simulations. Previous computational fluid dynamic (CFD) simulation work focused primarily on the pulsejet combustion and exhaust processes. This paper describes a new inlet sub-model which simulates the fluidic and mechanical operation of a valved pulsejet head. The governing equations for this sub-model are described. Sub-model validation is provided through comparisons of simulated and experimentally measured reed valve motion, and time averaged inlet mass flow rate. The updated pulsejet simulation, with the inlet sub-model implemented, is validated through comparison with experimentally measured combustion chamber pressure, inlet mass flow rate, operational frequency, and thrust. Additionally, the simulated pulsejet exhaust flowfield, which is dominated by a starting vortex ring, is compared with particle imaging velocimetry (PIV) measurements on the bases of velocity, vorticity, and vortex location. The results show good agreement between simulated and experimental data. The inlet sub-model is shown to be critical for the successful modeling of pulsejet operation. This sub-model correctly predicts both the inlet mass flow rate and its phase relationship with the combustion chamber pressure. As a result, the predicted pulsejet thrust agrees very well with experimental data.

Fully Integrated, Miniature, High-Frequency Flow Probe Utilizing MEMS Leadless SOI Technology

NASA Technical Reports Server (NTRS)

Ned, Alex; Kurtz, Anthony; Shang, Tonghuo; Goodman, Scott; Giemette. Gera (d)

2013-01-01

This work focused on developing, fabricating, and fully calibrating a flowangle probe for aeronautics research by utilizing the latest microelectromechanical systems (MEMS), leadless silicon on insulator (SOI) sensor technology. While the concept of angle probes is not new, traditional devices had been relatively large due to fabrication constraints; often too large to resolve flow structures necessary for modern aeropropulsion measurements such as inlet flow distortions and vortices, secondary flows, etc. Mea surements of this kind demanded a new approach to probe design to achieve sizes on the order of 0.1 in. (.3 mm) diameter or smaller, and capable of meeting demanding requirements for accuracy and ruggedness. This approach invoked the use of stateof- the-art processing techniques to install SOI sensor chips directly onto the probe body, thus eliminating redundancy in sensor packaging and probe installation that have historically forced larger probe size. This also facilitated a better thermal match between the chip and its mount, improving stability and accuracy. Further, the leadless sensor technology with which the SOI sensing element is fabricated allows direct mounting and electrical interconnecting of the sensor to the probe body. This leadless technology allowed a rugged wire-out approach that is performed at the sensor length scale, thus achieving substantial sensor size reductions. The technology is inherently capable of high-frequency and high-accuracy performance in high temperatures and harsh environments.

Experimental investigation of the radiation of sound from an unflanged duct and a bellmouth, including the flow effect

NASA Technical Reports Server (NTRS)

Ville, J. M.; Silcox, R. J.

1980-01-01

The radiation of sound from an inlet as a function of flow velocity, frequency, duct mode structure, and inlet geometry was examined by using a spinning mode synthesizer to insure a given space-time structure inside the duct. Measurements of the radiation pattern (amplitude and phase) and of the pressure reflection coefficient were obtained over an azimuthal wave number range of 0 to 6 and a frequency range up to 5000 Hz for an unflanged duct and a bellmouth. The measured radiated field and pressure reflection coefficient without flow for the unflanged duct agree reasonably well with theory. The influence of the inlet contour appears to be very drastic near the cut-on frequency of a mode and reasonable agreement is found between the bellmouth pressure reflection coefficient and a infinite hyperboloidal inlet theory. It is also shown that the flow has a weak effect on the amplitude of the directivity factor but significantly shifts the directivity factor phase. The influence of the flow on the modulus of the pressure reflection coefficient is found to be well described by a theoretical prediction.

System and method for continuous solids slurry depressurization

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

Leininger, Thomas Frederick; Steele, Raymond Douglas; Cordes, Stephen Michael

A system includes a first pump having a first outlet and a first inlet, and a controller. The first pump is configured to continuously receive a flow of a slurry into the first outlet at a first pressure and to continuously discharge the flow of the slurry from the first inlet at a second pressure less than the first pressure. The controller is configured to control a first speed of the first pump against the flow of the slurry based at least in part on the first pressure, wherein the first speed of the first pump is configured to resistmore » a backflow of the slurry from the first outlet to the first inlet.« less

Investigation of flow fields within large scale hypersonic inlet models

NASA Technical Reports Server (NTRS)

Gnos, A. V.; Watson, E. C.; Seebaugh, W. R.; Sanator, R. J.; Decarlo, J. P.

1973-01-01

Analytical and experimental investigations were conducted to determine the internal flow characteristics in model passages representative of hypersonic inlets for use at Mach numbers to about 12. The passages were large enough to permit measurements to be made in both the core flow and boundary layers. The analytical techniques for designing the internal contours and predicting the internal flow-field development accounted for coupling between the boundary layers and inviscid flow fields by means of a displacement-thickness correction. Three large-scale inlet models, each having a different internal compression ratio, were designed to provide high internal performance with an approximately uniform static-pressure distribution at the throat station. The models were tested in the Ames 3.5-Foot Hypersonic Wind Tunnel at a nominal free-stream Mach number of 7.4 and a unit free-stream Reynolds number of 8.86 X one million per meter.

Navier-Stokes simulation of external/internal transonic flow on the forebody/inlet of the AV-8B Harrier II

NASA Technical Reports Server (NTRS)

Mysko, Stephen J.; Chyu, Wei J.; Stortz, Michael W.; Chow, Chuen-Yen

1993-01-01

In this work, the computation of combined external/internal transonic flow on the complex forebody/inlet configuration of the AV-8B Harrier II is performed. The actual aircraft has been measured and its surface and surrounding domain, in which the fuselage and inlet have a common wall, have been described using structured grids. The 'thin-layer' Navier-Stokes equations were used to model the flow along with the Chimera embedded multi-block technique. A fully conservative, alternating direction implicit (ADI), approximately factored, partially fluxsplit algorithm was employed to perform the computation. Comparisons to some experimental wind tunnel data yielded good agreement for flow at zero incidence and angle of attack. The aim of this paper is to provide a methodology or computational tool for the numerical solution of complex external/internal flows.

Velocity and pressure measurements in guide vane clearance gap of a low specific speed Francis turbine

NASA Astrophysics Data System (ADS)

Thapa, B. S.; Dahlhaug, O. G.; Thapa, B.

2016-11-01

In Francis turbine, a small clearance gap between the guide vanes and the cover plates is usually required to pivot guide vanes as a part of governing system. Deflection of cover plates and erosion of mating surfaces causes this gap to increase from its design value. The clearance gap induces the secondary flow in the distributor system. This effects the main flow at the runner inlet, which causes losses in efficiency and instability. A guide vane cascade of a low specific speed Francis turbine has been developed for experimental investigations. The test setup is able to produce similar velocity distributions at the runner inlet as that of a reference prototype turbine. The setup is designed for particle image velocimetry (PIV) measurements from the position of stay vane outlet to the position of runner inlet. In this study, velocity and pressure measurements are conducted with 2 mm clearance gap on one side of guide vane. Leakage flow is observed and measured together with pressure measurements. It is concluded that the leakage flow behaves as a jet and mixes with the main flow in cross-wise direction and forms a vortex filament. This causes non-uniform inlet flow conditions at runner blades.

Aerodynamic Effects of Turbulence Intensity on a Variable-Speed Power-Turbine Blade with Large Incidence and Reynolds Number Variations

NASA Technical Reports Server (NTRS)

Flegel, Ashlie Brynn; Giel, Paul W.; Welch, Gerard E.

2014-01-01

The effects of inlet turbulence intensity on the aerodynamic performance of a variable speed power turbine blade are examined over large incidence and Reynolds number ranges. Both high and low turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow effects that are anticipated at cruise Reynolds numbers. The high turbulence study extends this to LPT-relevant turbulence levels while perhaps sacrificing transitional flow effects. Downstream total pressure and exit angle data were acquired for ten incidence angles ranging from +15.8 to 51.0. For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number ranging from 2.12105 to 2.12106 and at a design exit Mach number of 0.72. In order to achieve the lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of the blade row. The inlet turbulence levels ranged from 0.25 - 0.4 for the low Tu tests and 8- 15 for the high Tu study. Tu measurements were also made farther upstream so that turbulence decay rates could be calculated as needed for computational inlet boundary conditions. Downstream flow field measurements were obtained using a pneumatic five-hole pitchyaw probe located in a survey plane 7 axial chord aft of the blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired for each flow condition as well. The blade loading data show that the suction surface separation that was evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative incidence angles, the data show substantial differences in the exit flow field. These differences are attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the turbulence grid imposes.

Theoretical analysis of incompressible flow through a radial-inlet centrifugal impeller at various weight flows

NASA Technical Reports Server (NTRS)

Kramer, James J; Prian, Vasily D; Wu, Chung-Hua

1956-01-01

A method for the solution of the incompressible nonviscous flow through a centrifugal impeller, including the inlet region, is presented. Several numerical solutions are obtained for four weight flows through an impeller at one operating speed. These solutions are refined in the leading-edge region. The results are presented in a series of figures showing streamlines and relative velocity contours. A comparison is made with the results obtained by using a rapid approximate method of analysis.

Numerical simulation of flow and melting characteristics of seawater-ice crystals two-phase flow in inlet straight pipe of shell and tube heat exchanger of polar ship

NASA Astrophysics Data System (ADS)

Xu, Li; Huang, Chang-Xu; Huang, Zhen-Fei; Sun, Qiang; Li, Jie

2018-05-01

The ice crystal particles are easy to enter into the seawater cooling system of polar ship together with seawater when it sails in the Arctic. They are easy to accumulate in the pipeline, causing serious blockage of the cooling pipe. In this study, the flow and melting characteristics of ice particles-seawater two-phase flow in inlet straight pipe of shell-and-tube heat exchanger were numerically simulated by using Eulerian-Eulerian two-fluid model coupled with the interphase heat and mass transfer model. The influences of inlet ice packing factor, ice crystal particle diameter, and inlet velocity on the distribution and melting characteristics of ice crystals were investigated. The degree of asymmetry of the distribution of ice crystals in the cross section decreases gradually when the IPF changes from 5 to 15%. The volume fractions of ice crystals near the top of the outlet cross section are 19.59, 19.51, and 22.24% respectively for ice packing factor of 5, 10 and 15%. When the particle diameter is 0.5 mm, the ice crystals are gradually stratified during the flow process. With particle diameters of 1.0 and 2.0 mm, the region with the highest volume fraction of ice crystals is a small circle and the contours in the cloud map are compact. The greater the inlet flow velocity, the less stratified the ice crystals and the more obvious the turbulence on the outlet cross section. The average volume fraction of ice crystals along the flow direction is firstly rapidly reduced and then stabilized after 300 mm.

Comparison of numerical and experimental results of the flow in the U9 Kaplan turbine model

NASA Astrophysics Data System (ADS)

Petit, O.; Mulu, B.; Nilsson, H.; Cervantes, M.

2010-08-01

The present work compares simulations made using the OpenFOAM CFD code with experimental measurements of the flow in the U9 Kaplan turbine model. Comparisons of the velocity profiles in the spiral casing and in the draft tube are presented. The U9 Kaplan turbine prototype located in Porjus and its model, located in Älvkarleby, Sweden, have curved inlet pipes that lead the flow to the spiral casing. Nowadays, this curved pipe and its effect on the flow in the turbine is not taken into account when numerical simulations are performed at design stage. To study the impact of the inlet pipe curvature on the flow in the turbine, and to get a better overview of the flow of the whole system, measurements were made on the 1:3.1 model of the U9 turbine. Previously published measurements were taken at the inlet of the spiral casing and just before the guide vanes, using the laser Doppler anemometry (LDA) technique. In the draft tube, a number of velocity profiles were measured using the LDA techniques. The present work extends the experimental investigation with a horizontal section at the inlet of the draft tube. The experimental results are used to specify the inlet boundary condition for the numerical simulations in the draft tube, and to validate the computational results in both the spiral casing and the draft tube. The numerical simulations were realized using the standard k-e model and a block-structured hexahedral wall function mesh.

Size-selective sampling performance of six low-volume “total” suspended particulate (TSP) inlets

EPA Science Inventory

Several low-volume inlets (flow rates ≤ 16.7 liters per minute (Lpm)) are commercially available as components of low-cost, portable ambient particulate matter samplers. Because the inlets themselves do not contain internal fractionators, they are often assumed to representati...

Numerical simulation of supersonic inlets using a three-dimensional viscous flow analysis

NASA Technical Reports Server (NTRS)

Anderson, B. H.; Towne, C. E.

1980-01-01

A three dimensional fully viscous computer analysis was evaluated to determine its usefulness in the design of supersonic inlets. This procedure takes advantage of physical approximations to limit the high computer time and storage associated with complete Navier-Stokes solutions. Computed results are presented for a Mach 3.0 supersonic inlet with bleed and a Mach 7.4 hypersonic inlet. Good agreement was obtained between theory and data for both inlets. Results of a mesh sensitivity study are also shown.

Prediction of Laminar and Turbulent Boundary Layer Flow Separation in V/STOL Engine Inlets

NASA Technical Reports Server (NTRS)

Chou, D. C.; Luidens, R. W.; Stockman, N. O.

1977-01-01

A description is presented of the development of the boundary layer on the lip and diffuser surface of a subsonic inlet at arbitrary operating conditions of mass flow rate, free stream velocity and incidence angle. Both laminar separation on the lip and turbulent separation in the diffuser are discussed. The agreement of the theoretical results with model experimental data illustrates the capability of the theory to predict separation. The effects of throat Mach number, inlet size, and surface roughness on boundary layer development and separation are illustrated.

Improved computer programs for calculating potential flow in propulsion system inlets

NASA Technical Reports Server (NTRS)

Stockman, N. O.; Farrell, C. A., Jr.

1977-01-01

Computer programs to calculate the incompressible potential flow corrected for compressibility in axisymmetric inlets at arbitrary operating conditions are presented. Included are a statement of the problem to be solved, a description of each of the programs and sufficient documentation, including a test case, to enable a user to run the programs.

STUDY PROGRAM FOR TURBO-COOLER FOR PRODUCING ENGINE COOLING AIR.

DTIC Science & Technology

VANES , STAGNATION POINT, DECELERATION, ACCELERATION, SUPERSONIC DIFFUSERS, TURBINE BLADES , EVAPOTRANSPIRATION, LIQUID COOLED, HEAT TRANSFER, GAS BEARINGS, SEALS...HYPERSONIC AIRCRAFT , COOLING + VENTILATING EQUIPMENT), (*GAS TURBINES , COOLING + VENTILATING EQUIPMENT), HYPERSONIC FLOW, AIR COOLED, AIRCRAFT ... ENGINES , FEASIBILITY STUDIES, PRESSURE, SUPERSONIC CHARACTERISTICS, DESIGN, HEAT EXCHANGERS, COOLING (U) AXIAL FLOW TURBINES , DUCT INLETS, INLET GUIDE

40 CFR 60.54 - Test methods and procedures.

Code of Federal Regulations, 2011 CFR

2011-07-01

... sample CO2 concentrations at all traverse points. (ii) If sampling is conducted after a wet scrubber, an... volumetric flow rates at the inlet and outlet of the wet scrubber and the inlet CO2 concentration may be used... concentration measured before the scrubber, percent dry basis. Qdi=volumetric flow rate of effluent gas before...

Internal aerodynamics of a generic three-dimensional scramjet inlet at Mach 10

NASA Technical Reports Server (NTRS)

Holland, Scott D.

1995-01-01

A combined computational and experimental parametric study of the internal aerodynamics of a generic three-dimensional sidewall compression scramjet inlet configuration at Mach 10 has been performed. The study was designed to demonstrate the utility of computational fluid dynamics as a design tool in hypersonic inlet flow fields, to provide a detailed account of the nature and structure of the internal flow interactions, and to provide a comprehensive surface property and flow field database to determine the effects of contraction ratio, cowl position, and Reynolds number on the performance of a hypersonic scramjet inlet configuration. The work proceeded in several phases: the initial inviscid assessment of the internal shock structure, the preliminary computational parametric study, the coupling of the optimized configuration with the physical limitations of the facility, the wind tunnel blockage assessment, and the computational and experimental parametric study of the final configuration. Good agreement between computation and experimentation was observed in the magnitude and location of the interactions, particularly for weakly interacting flow fields. Large-scale forward separations resulted when the interaction strength was increased by increasing the contraction ratio or decreasing the Reynolds number.

Preliminary Results of an Altitude-Wind-Tunnel Investigation of an Axial-Flow Gas Turbine-Propeller Engine. 4; Compressor and Turbine Performance Characteristics

NASA Technical Reports Server (NTRS)

Wallner, Lewis E.; Saari, Martin J.

1948-01-01

As part of an investigation of the performance and operational characteristics of the axial-flow gas turbine-propeller engine, conducted in the Cleveland altitude wind tunnel, the performance characteristics of the compressor and the turbine were obtained. The data presented were obtained at a compressor-inlet ram-pressure ratio of 1.00 for altitudes from 5000 to 35,000 feet, engine speeds from 8000 to 13,000 rpm, and turbine-inlet temperatures from 1400 to 2100 R. The highest compressor pressure ratio obtained was 6.15 at a corrected air flow of 23.7 pounds per second and a corrected turbine-inlet temperature of 2475 R. Peak adiabatic compressor efficiencies of about 77 percent were obtained near the value of corrected air flow corresponding to a corrected engine speed of 13,000 rpm. This maximum efficiency may be somewhat low, however, because of dirt accumulations on the compressor blades. A maximum adiabatic turbine efficiency of 81.5 percent was obtained at rated engine speed for all altitudes and turbine-inlet temperatures investigated.

A CFD Case Study of a Fan Stage with Split Flow Path Subject to Total Pressure Distortion Inflow

NASA Technical Reports Server (NTRS)

To, Wai-Ming

2017-01-01

This report is the documentation of the work performed under the Hypersonic Project of the NASA's Fundamental Aeronautics Program. It was funded through Task Number NNC10E444T under GESS-2 Contract NNC06BA07B. The objective of the task is to develop advanced computational tools for the simulation of multi-stage turbomachinery in support of aeropropulsion. This includes work elements in extending the TURBO code and validating the multi-stage URANS (Unsteady Reynolds Averaged Navier Stokes) simulation results with the experimental data. The unsteady CFD (Computation Fluid Dynamics) calculations were performed in full wheel mode with and without screen generated total pressure distortion at the computational inflow boundary, as well as in single passage phase lag mode for uniform inflow. The experimental data were provided by NASA from the single stage RTA (Revolutionary Turbine Accelerator) fan test program.Significant non-uniform flow condition at the fan-face of the aeropropulsion system is frequentlyencountered in many of the advanced aerospace vehicles. These propulsion systems can be eithera podded or an embedded design employed in HWB (Hybrid Wing Body) airframe concept. It isalso a topic of interest in military applications, in which advanced air vehicles have already deployedsome form of embedded propulsion systems in their design because of the requirementsof compact and low observable inlets. Even in the conventional airframe/engine design, the fancould operate under such condition when the air vehicle is undergoing rapid maneuvering action.It is believed that a better understanding of the fan’s aerodynamic and aeromechanical responseto this type of operating condition or off design operation would be beneficial to designing distortiontolerant blades for improved engine operability.The objective for this research is to assess the capability of turbomachinery code as an analysistool in understanding the effects and evaluating the impact of flow distortion on the aerodynamicand aeromechanical performance of the fan in advanced propulsion systems. Results from thetesting of an advanced fan stage released by NASA are available and will be used here for CFDcode validation. The experiment was performed at NASA’s high speed compressor facility aspart of the RTA (Revolutionary Turbine Accelerator) demonstration project, a joint effort ofNASA Glenn Research Center and GE Aircraft Engines in developing an advanced Mach 4TBCC (Turbine Based Combined Cycle) turbofan/ramjet engine for access to space. Part of thetest was to assess the aerodynamic performance and operability of the fan stage under nonuniforminflow condition. Various flow distortion patterns were created at the fan-face by manipulatingsets of screens placed upstream of the wind tunnel. Measurements at the fan-face willprovide the necessary distortion flow information as the inflow boundary condition for the CFDin a full wheel simulation. Therefore the purpose of this work is to demonstrate the NASA supportedmulti-stage turbomachinery code, TURBO [1-5], in the aerodynamic performance analysisof a modern fan design operating under off design condition, and in particular to validate theCFD results with the RTA fan test data.A brief description of the RTA fan rig configuration is given in the next section, explaining onhow flow distortion were measured in the test and constructed for the CFD at the fan-face. It isfollowed by a section summarizing previous CFD work performed at NASA relevant to the currentfan configuration. A short description of the TURBO code is given next, followed by detailsin the computational model of the fan rig, the required computing resources, and the numericalprocedure for the simulations. The CFD results are presented in the discussion section and finallyconcluding remarks are summarized.

Characterization of flow in a scroll duct

NASA Technical Reports Server (NTRS)

Begg, E. K.; Bennett, J. C.

1985-01-01

A quantitative, flow visualization study was made of a partially elliptic cross section, inward curving duct (scroll duct), with an axial outflow through a vaneless annular cutlet. The working fluid was water, with a Re(d) of 40,000 at the inlet to the scroll duct, this Reynolds number being representative of the conditions in an actual gas turbine scroll. Both still and high speed moving pictures of fluorescein dye injected into the flow and illuminated by an argon ion laser were used to document the flow. Strong secondary flow, similar to the secondary flow in a pipe bend, was found in the bottom half of the scroll within the first 180 degs of turning. The pressure field set up by the turning duct was strong enough to affect the inlet flow condition. At 90 degs downstream, the large scale secondary flow was found to be oscillatory in nature. The exit flow was nonuniform in the annular exit. By 270 degs downstream, the flow appeared unorganized with no distinctive secondary flow pattern. Large scale structures from the upstream core region appeared by 90 degs and continued through the duct to reenter at the inlet section.

NASCRIN - NUMERICAL ANALYSIS OF SCRAMJET INLET

NASA Technical Reports Server (NTRS)

Kumar, A.

1994-01-01

The NASCRIN program was developed for analyzing two-dimensional flow fields in supersonic combustion ramjet (scramjet) inlets. NASCRIN solves the two-dimensional Euler or Navier-Stokes equations in conservative form by an unsplit, explicit, two-step finite-difference method. A more recent explicit-implicit, two-step scheme has also been incorporated in the code for viscous flow analysis. An algebraic, two-layer eddy-viscosity model is used for the turbulent flow calculations. NASCRIN can analyze both inviscid and viscous flows with no struts, one strut, or multiple struts embedded in the flow field. NASCRIN can be used in a quasi-three-dimensional sense for some scramjet inlets under certain simplifying assumptions. Although developed for supersonic internal flow, NASCRIN may be adapted to a variety of other flow problems. In particular, it should be readily adaptable to subsonic inflow with supersonic outflow, supersonic inflow with subsonic outflow, or fully subsonic flow. The NASCRIN program is available for batch execution on the CDC CYBER 203. The vectorized FORTRAN version was developed in 1983. NASCRIN has a central memory requirement of approximately 300K words for a grid size of about 3,000 points.

Nonlinear optimal control policies for buoyancy-driven flows in the built environment

NASA Astrophysics Data System (ADS)

Nabi, Saleh; Grover, Piyush; Caulfield, Colm

2017-11-01

We consider optimal control of turbulent buoyancy-driven flows in the built environment, focusing on a model test case of displacement ventilation with a time-varying heat source. The flow is modeled using the unsteady Reynolds-averaged equations (URANS). To understand the stratification dynamics better, we derive a low-order partial-mixing ODE model extending the buoyancy-driven emptying filling box problem to the case of where both the heat source and the (controlled) inlet flow are time-varying. In the limit of a single step-change in the heat source strength, our model is consistent with that of Bower et al.. Our model considers the dynamics of both `filling' and `intruding' added layers due to a time-varying source and inlet flow. A nonlinear direct-adjoint-looping optimal control formulation yields time-varying values of temperature and velocity of the inlet flow that lead to `optimal' time-averaged temperature relative to appropriate objective functionals in a region of interest.

ON THE MODIFICATION OF THE LOW FLOW-RATE PM10 DICHOTOMOUS SAMPLER INLET

EPA Science Inventory

A popular flat-topped inlet used for the collection of atmospheric particulate matter was modified to reduce water intrusion during rain and snow events. Simple alterations in the intake region of this inlet were made, including a larger drain hole, a one piece top plate, and ...

Staged membrane oxidation reactor system

DOEpatents

Repasky, John Michael; Carolan, Michael Francis; Stein, VanEric Edward; Chen, Christopher Ming-Poh

2014-05-20

Ion transport membrane oxidation system comprising (a) two or more membrane oxidation stages, each stage comprising a reactant zone, an oxidant zone, one or more ion transport membranes separating the reactant zone from the oxidant zone, a reactant gas inlet region, a reactant gas outlet region, an oxidant gas inlet region, and an oxidant gas outlet region; (b) an interstage reactant gas flow path disposed between each pair of membrane oxidation stages and adapted to place the reactant gas outlet region of a first stage of the pair in flow communication with the reactant gas inlet region of a second stage of the pair; and (c) one or more reactant interstage feed gas lines, each line being in flow communication with any interstage reactant gas flow path or with the reactant zone of any membrane oxidation stage receiving interstage reactant gas.

Staged membrane oxidation reactor system

DOEpatents

Repasky, John Michael; Carolan, Michael Francis; Stein, VanEric Edward; Chen, Christopher Ming-Poh

2013-04-16

Ion transport membrane oxidation system comprising (a) two or more membrane oxidation stages, each stage comprising a reactant zone, an oxidant zone, one or more ion transport membranes separating the reactant zone from the oxidant zone, a reactant gas inlet region, a reactant gas outlet region, an oxidant gas inlet region, and an oxidant gas outlet region; (b) an interstage reactant gas flow path disposed between each pair of membrane oxidation stages and adapted to place the reactant gas outlet region of a first stage of the pair in flow communication with the reactant gas inlet region of a second stage of the pair; and (c) one or more reactant interstage feed gas lines, each line being in flow communication with any interstage reactant gas flow path or with the reactant zone of any membrane oxidation stage receiving interstage reactant gas.

Staged membrane oxidation reactor system

DOEpatents

Repasky, John Michael; Carolan, Michael Francis; Stein, VanEric Edward; Chen, Christopher Ming-Poh

2012-09-11

Ion transport membrane oxidation system comprising (a) two or more membrane oxidation stages, each stage comprising a reactant zone, an oxidant zone, one or more ion transport membranes separating the reactant zone from the oxidant zone, a reactant gas inlet region, a reactant gas outlet region, an oxidant gas inlet region, and an oxidant gas outlet region; (b) an interstage reactant gas flow path disposed between each pair of membrane oxidation stages and adapted to place the reactant gas outlet region of a first stage of the pair in flow communication with the reactant gas inlet region of a second stage of the pair; and (c) one or more reactant interstage feed gas lines, each line being in flow communication with any interstage reactant gas flow path or with the reactant zone of any membrane oxidation stage receiving interstage reactant gas.

Mean flowfields in axisymmetric combustor geometries with swirl

NASA Astrophysics Data System (ADS)

Rhode, D. L.; Lilley, D. G.; McLaughlin, D. K.

1982-01-01

Six flowfield configurations are investigated with sidewall angles of 90 and 45 deg, and swirl vane angles of 0, 45, and 70 deg. It is found that central recirculation zones occur for the swirling flow cases investigated, which extend from the inlet to x/D = 1.7, where x is the axial polar coordinate, and D is the test section diameter. Five-hole pitot probe pressure measurements are used to determine time-mean velocities, and corresponding flow situations are predicted and compared to results of experimental data. Excellent agreement is found for the nonswirling flow, although poor agreement is found for swirling flow cases, especially near the inlet. The discrepancy is attributed to the lack of realism in the turbulence model, and/or to inaccurate specification of time-mean velocity and turbulence energy distributions at the inlet.

Low-speed performance of an axisymmetric, mixed-compression, supersonic inlet with auxiliary inlets

NASA Technical Reports Server (NTRS)

Trefny, C. J.; Wasserbauer, J. W.

1986-01-01

A test program was conducted to determine the aerodynamic performance and acoustic characteristics associated with the low-speed operation of a supersonic, axisymmetric, mixed-compression inlet with auxiliary inlets. Blow-in-auxiliary doors were installed on the NASA Ames P inlet. One door per quadrant was located on the cowl in the subsonic diffuser selection of the inlet. Auxiliary inlets with areas of 20 and 40 percent of the inlet capture area were tested statically and at free-stream Mach numbers of 0.1 and 0.2. The effects of boundary layer bleed inflow were investigated. A JT8D fan simulator driven by compressed air was used to pump inlet flow and to provide a characteristic noise signature. Baseline data were obtained at static free-stream conditions with the sharp P-inlet cowl lip replaced by a blunt lip. Auxiliary inlets increased overall total pressure recovery of the order of 10 percent.