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

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At least 415 records · Page 23
All sourcesNASA NTRSDOE OSTI
NASA NTRS1946

Effects of Wing Flaps and Wing Duct Inlet on the Lift and Stalling Characteristics of a 1/4-Scale Partial-Span Model of the Republic XF-12 Airplane in the Langley 19-Foot Pressure Tunnel

An investigation was conducted in the Langley 19-foot pressure tunnel to determine the lift, drag, pitching-moment and stalling characteristics fo a 1/4 -scale partial-span model of the left wing of the Republic XF-12 airplane. The effects of a duct inlet, located between the nacelles at the leading edge of the wing, on those characteristics were also investigated. The Reynolds numbers for the investigation covered a range from 4,500,000 to 8,600,000. The results of the investigation indicated that maximum lift coefficients of 1.36, 1.71, and 2.11 were measured on the model with flaps neutral and deflected 20 deg and 55 deg, respectively at a Reynolds number of 8,600,000. When the duct inlet was replaced by a basic airfoil nose the flap-neutral maximum-lift coefficient was increased from 1.36 to 1.41. The results also showed that at maximum lift with flaps neutral or deflected 55 deg. most of the area between the nacelles were stalled while only small areas on other portions of the model were stalled; when the duct inlet was replaced by the basic airfoil nose the stall was delayed to a slightly higher angle of attack but the nature of the stall was relatively unaffected.

Graham, Robert R.↗
NASA NTRS1951

Low-Lift Drag and Duct Pressure Recovery of a 1/8.25-Scale Model of the Consolidated Vultee XF-92 Airplane at Mach Numbers from 0.7 to 1.4

A flight investigation has been made to determine the external drag and pressure recovery of a 1/8.25 - scale flight model of the Consolidated Vultee XF-92 from Mach numbers 0.7 to 1.4 and Reynolds numbers from 8.5 x 10(exp 6) to 19.2 x 10(exp 6) at or near zero lift. Relative mass flow, average pressure recovery, total drag, internal drag, and external drag are presented as functions of Mach number. Between Mach numbers of 0.90 and 0.975, the external drag of the configuration (including base drag of the inner body and additive drag) was about equal to that of a similar model with a faired nose and no mass flow; however, at supersonic speeds the drag coefficient for the faired-nose model remained relatively constant whereas the drag coefficient for the ducted model continued to increase sharply. The internal drag coefficient of the duct was roughly constant at 0.013 up to a Mach number of 1.20; after which it decreased to 0.0075 at a Mach number of 1.4. The over-all pressure recovery of the inlet and duct varied from 94 percent at a Mach number of 0.7 to about 91 percent at a Mach number of 1.4 at a relative-mass-flow ratio of about 0.30. The losses in pressure recovery were believed to be caused by the possible occurrence of separation of flow from the inner body and by an aerodynamically unclean internal configuration which did not duplicate the form proposed for the original XF-92 airplane.

Mitcham, Grady L.↗
NASA NTRS2006

Rotating Rake Mode Measurements Over Passive Treatment in a Ducted Fan

The NASA Glenn Research Center s Rotating Rake mode measurement system has been successful in measuring the modal content propagating in hardwall ducts. This paper proposes an extension of the Rotating Rake measurement and analysis technique to treated sections by developing basis functions based on wall impedance boundary conditions for flow conditions (i.e., constant duct area and Mach number) where the closed form analytical solution exists. Analytical equations developed to estimate mode power are incorporated. This method is verified by decomposing and analyzing radial pressure profiles generated numerically by the Eversman propagation code. Several modes, frequencies and impedances are evaluated. Data from a low-speed ducted fan with several different impedance conditions was acquired and reduced to determine the best fit to the data. Using the impedance boundary conditions result in better mode measurement solutions.

Sutliff, Daniel L.↗
NASA NTRS1994

Ramjet bypass duct and preburner configuration

A combined turbofan and ramjet aircraft engine includes a forward bypass duct which allows the engine to operate more efficiently during the turbofan mode of operation. By mounting a ramjet preburner in the forward duct and isolating this duct from the turbofan bypass air, a transition from turbofan operation to ramjet operation can take place at lower flight Mach numbers without incurring pressure losses or blockage in the turbofan bypass air.

Orlando, Robert J.↗
NASA NTRS2008

Assessment of 3D Codes for Predicting Liner Attenuation in Flow Ducts

This paper presents comparisons of seven propagation codes for predicting liner attenuation in ducts with flow. The selected codes span the spectrum of methods available (finite element, parabolic approximation, and pseudo-time domain) and are collectively representative of the state-of-art in the liner industry. These codes are included because they have two-dimensional and three-dimensional versions and can be exported to NASA's Columbia Supercomputer. The basic assumptions, governing differential equations, boundary conditions, and numerical methods underlying each code are briefly reviewed and an assessment is performed based on two predefined metrics. The two metrics used in the assessment are the accuracy of the predicted attenuation and the amount of wall clock time to predict the attenuation. The assessment is performed over a range of frequencies, mean flow rates, and grazing flow liner impedances commonly used in the liner industry. The primary conclusions of the study are (1) predicted attenuations are in good agreement for rigid wall ducts, (2) the majority of codes compare well to each other and to approximate results from mode theory for soft wall ducts, (3) most codes compare well to measured data on a statistical basis, (4) only the finite element codes with cubic Hermite polynomials capture extremely large attenuations, and (5) wall clock time increases by an order of magnitude or more are observed for a three-dimensional code relative to the corresponding two-dimensional version of the same code.

Watson, W. R.↗
NASA NTRS1945

An Experimental Investigation of NACA Submerged-Duct Entrances

The results of an investigation of submerged-duct entrances are presented. It is shown that this type of entrance possesses the following characteristics: 1) very high-critical-compressibility speeds throughout the range of high-speed inlet velocity ratios; 2) very low pressure losses for the air entering the duct at all inlet-velocity ratios; and, 3) low external drag. These characteristics are obtained by the proper shaping of the contour of the upstream approach to the submerged inlets and by proper alignment of the duct lip. Design data are presented and the application of these data to a specific high-speed fighter-airplane design is discussed.

Frick, Charles W.↗
NASA NTRS2014

Computational Aerodynamic Simulations of an 840 ft/sec Tip Speed Advanced Ducted Propulsor Fan System Model for Acoustic Methods Assessment and Development

Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted Propulsor fan system were performed at five different operating points on the fan operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15- foot Low Speed Wind Tunnel at the NASA Glenn Research Center, resulting in quality, detailed aerodynamic and acoustic measurement data. Details of the fan geometry, the computational fluid dynamics methods, the computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating conditions simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system, excluding a long core duct section downstream of the core inlet guide vane. As a result, only fan rotational speed and system bypass ratio, set by specifying static pressure downstream of the core inlet guide vane row, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. The computed blade row flow fields for all five fan operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles, chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no excessive boundary layer separations or related secondary-flow problems. A few spanwise comparisons between computational and measurement data in the bypass duct show that they are in good agreement, thus providing a partial validation of the computational results.

Tweedt, Daniel L.↗
NASA NTRS2014

Reverse Flow Engine Core Having a Ducted Fan with Integrated Secondary Flow Blades

Secondary air flow is provided for a ducted fan having a reverse flow turbine engine core driving a fan blisk. The fan blisk incorporates a set of thrust fan blades extending from an outer hub and a set of integral secondary flow blades extending intermediate an inner hub and the outer hub. A nacelle provides an outer flow duct for the thrust fan blades and a secondary flow duct carries flow from the integral secondary flow blades as cooling air for components of the reverse flow turbine engine.

Kisska, Michael K.↗
NASA NTRS2018

Two-Dimensional Modal Beamforming in Wavenumber Space for Duct Acoustics

As turbofan bypass ratio continues to increase, civilian aircraft noise is increasingly dominated by fan noise. Fan noise propagating from its rotor and stator origins to the community passes through the inlet or aft flow duct, where its confined situation makes it susceptible to characterization by wall-mounted microphone arrays. Recently, the NASA-Glenn Research Center adapted its W-8 Single Stage Axial Compressor Facility to this type of measurement. OptiNav, Inc. took the opportunity to improve and simplify the duct mode processing in its Beamform Interactive computer program. A new approach to in-duct beamforming with a 2D wall-mounted array of microphones was developed. The purpose of this paper is to document the beamforming approach and provide some sample results from the W-8 facility.

Turbofan↗
NASA NTRSDate not supplied

Bypass Duct Acoustic Liner Design With and Without Bifurcation Effects

The growth in air traffic and the commitment to sustainable aviation continue to provide new challenges to reducing aircraft noise levels. Acoustic liner design methodologies must therefore provide the capability to efficiently predict the acoustic benefits of novel liner configurations within complex aircraft nacelle geometries. With these observations in mind, a broadband acoustic liner optimization process has been developed and assessed through a series of design and experimental studies at increasing technology readiness levels. This work applies the design process to the aft-fan noise component and explores the effects of bypass duct bifurcations (e.g., the pylon and lower bifurcation). In addition to this new application, the design study is expanded to include the use of a commercially available duct propagation code. Despite the different general workflow for the two propagation codes, consistent optimized impedance spectra and in-duct attenuation predictions were obtained for several acoustic treatment options. The preliminary results are promising, and this work increases confidence in the enhanced broadband liner design methodology and lays the groundwork for complimentary use of the codes in future studies. The potential benefits of acoustic treatment on the upper and lower bifurcations are also demonstrated. The knowledge gained through this preliminary stage of the liner design process will be used to guide the identification of candidate liner designs for a proposed static engine test within the next year.

Acoustic Liner Design↗
NASA NTRSDate not supplied

Bypass Duct Acoustic Liner Design with and without Bifurcation Effects

The growth in air traffic and the commitment to sustainable aviation continue to provide new challenges to reducing aircraft noise levels. Acoustic liner design methodologies must therefore provide the capability to efficiently predict the acoustic benefits of novel liner configurations within complex aircraft nacelle geometries. With these observations in mind, a broadband acoustic liner optimization process has been developed and assessed through a series of design and experimental studies at increasing technology readiness levels. This work applies the design process to the aft-fan noise component and explores the effects of bypass duct bifurcations (e.g., the pylon and lower bifurcation). In addition to this new application, the design study is expanded to include the use of a commercially available duct propagation code. Despite the different general workflow for the two propagation codes, consistent optimized impedance spectra and in-duct attenuation predictions were obtained for several acoustic treatment options. The preliminary results are promising, and this work increases confidence in the enhanced broadband liner design methodology and lays the groundwork for complimentary use of the codes in future studies. The potential benefits of acoustic treatment on the upper and lower bifurcations are also demonstrated. The knowledge gained through this preliminary stage of the liner design process will be used to guide the identification of candidate liner designs for a proposed static engine test within the next year.

Acoustic Liner Design↗
DOE OSTI2023

A computational parametric study of ducted fuel injection implementation in a heavy-duty diesel engine

Experiments have shown that ducted fuel injection (DFI) effectively reduces soot emissions from direct-injection diesel engines. Although many computational studies have evaluated DFI’s spray development and soot reduction mechanisms in constant volume chambers, only limited computational work on internal combustion engines exists. The DFI duct assembly changes the engine’s in-cylinder flow, spray, and combustion development. Therefore, current production engine designs might not be optimal for achieving the best engine performance with DFI. Here, this work conducted an extensive numerical study to evaluate how parameter changes affect DFI performance. The parameters include swirl ratio, piston geometry, compression ratio (CR), number of injector orifices, split injection strategy, and exhaust gas recirculation (EGR) in a heavy-duty diesel engine utilizing DFI. The combustion and soot emission data from the Sandia compression ignition optical research engine were used for model validation. Simulations showed that an increased swirl ratio resulted in more intense jet flame-piston interaction, slowing down the combustion heat release during the late combustion stage and leading to lower indicated thermal efficiency (ITE) due to higher exhaust losses. A piston-bowl design with a reentrant inner piston edge yielded the highest thermal efficiency, due to the reduced cylinder head heat transfer loss. Additional injector orifices led to higher efficiency owing to a more advanced combustion phasing. Nevertheless, the maximum pressure rise rate (MPRR) and oxides of nitrogen (NO x ) emissions also increased with the number of injector orifices due to more rapid heat release and higher combustion temperature. Implementation of a split injection strategy combined with a higher EGR rate effectively inhibited the excessive MPRR and NO x formation. In general, the study concluded that DFI is not sensitive to most parameter changes but will benefit from future parameter optimization.

33 ADVANCED PROPULSION SYSTEMS↗
DOE OSTI2023

Flow friction in a wide rectangular duct with a beveled inlet: Informal Notes of Fundamental Fluid Physics Studies

Fundamental streamwise pressure distributions were measured at Idaho State University to address a primary objective of obtaining insight into the apparent evolution of air flow regimes at moderate Reynolds numbers in thin, wide rectangular ducts with converging inlets. Ducts of 3.030 and 4.077 mm spacings s (called three and four mm), having non-dimensional lengths of (L/s) ≈ 500.8 and 372.2 and widths (W/s) ≈ 114.6 and 83.2, respectively, were compared. Flow rates ranged from Reynolds numbers based on hydraulic diameter (Re Dh ) of about 980 to 10,000 and corresponding non-dimensional downstream pressure gradient parameters (-K p ) covered the range 0.038 to 0.0061, respectively.

77 NANOSCIENCE AND NANOTECHNOLOGY↗
NASA NTRS1948

Ice Protection of Turbojet Engines by Inertia Separation of Water II : Single-offset-duct System

Investigation of a single-offset-duct system designed to prevent entrance of water into a turbojet engine was conducted on a half-scale nacelle model. An investigation was made to determine ram-pressure recovery and radial velocity profiles at the compressor section and icing characteristics of such a duct system. At a design inlet velocity of 0.77, the maximum ram-pressure recovery attained with effective water-separating inlet was 77 percent, which is considerably less than attainable with a direct-ram inlet. Continuous heating of the accessory-housing surface would be required for inlets that have a small ice storage space.

NOSE SHAPE - DUCTED BODIES↗
NASA NTRS2014

In-Duct and Far-Field Experimental Measrements from the ANCF for the Purpose of Improved Broadband Liner Optimization

A process for the design and evaluation of novel broadband acoustic liner concepts with limited fan source information is being evaluated. A pair of advanced broad-bandwidth liners were designed and manufactured for the NASA Glenn Research Center's Advanced Noise Control Fan (ANCF): (i) a constant impedance liner and (ii) a variable impedance liner. The insertion loss of both liners was measured in-duct utilizing the ANCF's Configurable Fan Artificial Noise System in a clean configuration with no-flow. Additionally, the acoustic characteristics of the Variable Impedance Liner were measured in the standard ANCF configuration with and without flow. The experimental setup, in-duct mode power levels, and far-field directivity are presented herein.

Acoustics↗
NASA NTRS2015

Numerical Predictions of Mode Reflections in an Open Circular Duct: Comparison with Theory

The NASA Broadband Aeroacoustic Stator Simulation code was used to compute the acoustic field for higher-order modes in a circular duct geometry. To test the accuracy of the results computed by the code, the duct was terminated by an open end with an infinite flange or no flange. Both open end conditions have a theoretical solution that was used to compare with the computed results. Excellent comparison for reflection matrix values was achieved after suitable refinement of the grid at the open end. The study also revealed issues with the level of the mode amplitude introduced into the acoustic held from the source boundary and the amount of reflection that occurred at the source boundary when a general nonreflecting boundary condition was applied.

Duct Modes↗
NASA NTRSDate not supplied

NASA Small Engine Components Compressor Test Facility: High Efficiency Centrifugal Compressor Vaneless Diffuser and Transition Duct Configurations

The original vaned configuration of the High Efficiency Centrifugal Compressor (HECC) was developed with aggressive design targets to facilitate advancement of state-of-the-art technology in gas turbine compressors. This work documents modifications to the Small Engine Components Compressor Test Facility to support testing of the HECC vaneless diffuser and transition duct configurations, both of which are open geometries. These configurations were developed as follow-up studies to improve understanding of the differences between the predicted and actual performance of the original HECC vaned diffuser stage. Specifically, the vaneless diffuser was designed to provide the impeller with ideal exit conditions such that the impeller performance could be isolated from downstream stationary components, and the additively manufactured transition duct inlet configuration enables investigation into the effects of the inlet flow path on the stage performance. Features of note in the facility are active clearance control, bleed flow regulation, modular inlet configurations, and steady-state and fast-response instrumentation throughout the flow path. Details of the facility and instrumentation are presented as well as the baseline performance of the compressor stage. The geometry, performance, and detailed aerodynamic data have been made available to the public at https://storage.googleapis.com/heccdata/NASA-HECC-Data-Archive.zip.

centrifugal compressor↗
NASA NTRSDate not supplied

NASA Small Engine Components Compressor Test Facility: High Efficiency Centrifugal Compressor Vaneless Diffuser and Transition Duct Configurations

The original vaned configuration of the High Efficiency Centrifugal Compressor (HECC) was developed with aggressive design targets to facilitate advancement of state-of-the-art technology in gas turbine compressors. This work documents modifications to the Small Engine Components Compressor Test Facility to support testing of the HECC vaneless diffuser and transition duct configurations, both of which are open geometries. These configurations were developed as follow-up studies to improve understanding of the differences between the predicted and actual performance of the original HECC vaned diffuser stage. Specifically, the vaneless diffuser was designed to provide the impeller with ideal exit conditions such that the impeller performance could be isolated from downstream stationary components, and the additively manufactured transition duct inlet configuration enables investigation into the effects of the inlet flow path on the stage performance. Features of note in the facility are active clearance control, bleed flow regulation, modular inlet configurations, and steady-state and fast-response instrumentation throughout the flow path. Details of the facility and instrumentation are presented as well as the baseline performance of the compressor stage. The geometry, performance, and detailed aerodynamic data have been made available to the public at https://storage.googleapis.com/hecc-data/NASA-HECC-Data-Archive.zip.

centrifugal compressor↗