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NASA NTRS1982

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

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

Vadyak, J.↗

NASA NTRS1972

Forward flight effects on mixer nozzle design and noise considerations for STOL externally blown flap systems

Experimental data of the peak axial-velocity decay in a moving airstream are presented for several types of nozzles. The nozzles include a six-tube mixer nozzle of a type considered for reduction of jet-flap interaction noise for externally-blown-flap STOL aircraft. The effect of secondary flow on the core flow velocity decay of a bypass nozzle is also discussed. Tentative correlation equations are suggested for the configurations evaluated. Recommendations for minimizing forward velocity effects on velocity decay and jet-flap interaction noise are made.

Vonglahn, U. H.↗

NASA NTRS1972

Very high temperature laminar flow of a gas through the entrance region of a cooled tube - Numerical calculations and experimental results.

The laminar flow equations in differential form are solved numerically on a digital computer for flow of a very high temperature gas through the entrance region of an externally cooled tube. The solution method is described and calculations are carried out in conjunction with experimental measurements. The agreement with experiment is good, with the result indicating relatively large energy and momentum losses in the highly cooled flows considered where the pressure is nearly uniform along the flow and the core flow becomes non-adiabatic a few diameters downstream of the inlet. The effects of a large range of Reynolds number and Mach number (viscous dissipation) are also investigated.

Back, L. H.↗

NASA NTRS1972

Forward flight effects on mixer nozzle design and noise considerations for STOL externally blown flap systems.

Vonglahn, U. H.↗

NASA NTRS1984

Theoretical approach to labyrinth seal forces - cross-coupled stiffness of a straight-through labyrinth seal

Two kinds of three dimensional flows in a labyrinth seal, a jet flow and a core flow, are considered and theoretical equations are set up concerning the motion of each flow. The pressure distribution within the labyrinth is calculated, when the rotor shaft makes a small displacement from the center line of the casing, keeping parallel with it. The theoretical values of cross coupled stiffness obtained by integrating the pressure under different labyrinth geometries and operating conditions through these formulas are compared with the experimental data.

Kameoka, T.↗

NASA NTRS1994

Optimization of Orifice Geometry for Cross-Flow Mixing in a Cylindrical Duct

Mixing of gaseous jets in a cross-flow has significant applications in engineering, one example of which is the dilution zone of a gas turbine combustor. Despite years of study, the design of jet injection in combustors is largely based on practical experience. A series of experiments was undertaken to delineate the optimal mixer orifice geometry. A cross-flow to core-flow momentum-flux ratio of 40 and a mass flow ratio of 2.5 were selected as representative of an advanced design. An experimental test matrix was designed around three variables: the number of orifices, the orifice aspect ratio (long-to-short dimension), and the orifice angle. A regression analysis was performed on the data to arrive at an interpolating equation that predicted the mixing performance of orifice geometry combinations within the range of the test matrix parameters. Results indicate that mixture uniformity is a non-linear function of the number of orifices, the orifice aspect ratio, and the orifice angle. Optimum mixing occurs when the asymptotic mean jet trajectories are in the range of 0.35 less than r/R less than 0.5 (where r = 0 is at the mixer wall) at z/R = 1.0. At the optimum number of orifices, the difference between shallow-angled slots with large aspect ratios and round holes is minimal and either approach will lead to good mixing performance. At the optimum number of orifices, it appears possible to have two local optimums where one corresponds to an aspect ratio of 1.0 and the other to a high aspect ratio.

Sowa, W. A.↗

NASA NTRS1996

Optimization of Orifice Geometry for Cross-Flow Mixing in a Cylindrical Duct

Mixing of gaseous jets in a cross-flow has significant applications in engineering, one example of which is the dilution zone of a gas turbine combustor. Despite years of study, the design of the jet injection in combustors is largely based on practical experience. The emergence of NO(x) regulations for stationary gas turbines and the anticipation of aero-engine regulations requires an improved understanding of jet mixing as new combustor concepts are introduced. For example, the success of the staged combustor to reduce the emission of NO(x) is almost entirely dependent upon the rapid and complete dilution of the rich zone products within the mixing section. It is these mixing challenges to which the present study is directed. A series of experiments was undertaken to delineate the optimal mixer orifice geometry. A cross-flow to core-flow momentum-flux ratio of 40 and a mass flow ratio of 2.5 were selected as representative of a conventional design. An experimental test matrix was designed around three variables: the number of orifices, the orifice length-to- width ratio, and the orifice angle. A regression analysis was performed on the data to arrive at an interpolating equation that predicted the mixing performance of orifice geometry combinations within the range of the test matrix parameters. Results indicate that the best mixing orifice geometry tested involves eight orifices with a long-to-short side aspect ratio of 3.5 at a twenty-three degree inclination from the center-line of the mixing section.

Kroll, J. T.↗

NASA NTRSDate not supplied

Acoustic Characterization of the NASA Langley 14- by 22-Foot Subsonic Tunnel Using Single-Microphone Analysis Techniques

An experimental campaign was conducted to assess recent acoustic modifications to the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel. This effort was undertaken in preparation for future rotorcraft, Advanced Air Mobility, and airframe noise acoustic tests. The tunnel is a closed circuit and typically operates in an open-jet configuration for acoustic studies. A vertical linear array of microphones and a phased array were placed on traverses outside of the core flow. Pole-mounted acoustic sources with known waveforms were used to identify reflective surfaces under static conditions with no tunnel flow. A similar process was replicated with more compact sources in an aerodynamic fairing for flow speeds up to Mach 0.16. This enabled investigation of the test section core flow and bounding shear layer impacts on the acoustic measurements. Periodic averaging was employed and was shown to be capable of isolating periodic acoustic signals even for poor signal-to-noise ratio conditions. Benefits and limitations of single-microphone processing methods are identified. A companion paper utilizes a phased array in an effort to address the identified limitations to more traditional single-microphone data collection.

Acoustic Characterization↗

NASA NTRSDate not supplied

Acoustic Characterization of the NASA Langley 14- by 22-Foot Subsonic Tunnel Using Single-Microphone Analysis Techniques

Acoustic Characterization↗

NASA NTRS2020

1- by 1-Foot Supersonic Wind Tunnel Sidewall Flowfield Compression Plate Study

A method for locally increasing the dynamic pressure in the test section of a supersonic wind tunnel is investigated from a theoretical, computational and experimental perspective, culminating in a reduced-scale, proof-of-concept test at Mach 3.44. Symmetric, variable-angle compression plates along the test section sidewalls generate oblique shockwaves to elevate the test section dynamic pressure (q). The experimental effort utilizes boundary-layer pitot rakes to sample the incoming flow near the floor, pitot and static rakes to measure the high-q core flow and numerous surface taps to map out the static pressure distribution along the floor, ceiling and sidewalls. A reliable and robust starting procedure is developed to establish supersonic flow in the test section and behind the compression plates. A characterization of the incoming floor boundary layer reveals previously documented features (such as a bulge toward the centerline) and an unexpected (and as-yet-unexplained) thickening of the boundary layer with increasing Reynolds number around total pressures of 40–60 psia. The high-q core flow is characterized for eight compression plate angles (4.0–8.0°), eight tunnel total pressures (20–139 psia) and three axial stations (24.95–27.35 inches). The peak dynamic pressure rise exceeds expectations. At conditions relevant to full-scale testing (6.0° and 139 psia), the dynamic pressure increases by a factor of 1.73, the Mach number becomes 2.85 and the total pressure drops imperceptibly. Basic distortion metrics show transverse non-uniformity of less than 5 %. Limited spatial resolution in the static rake measurements warrants additional, higher fidelity testing or computational analysis. Surface-mounted, fast-response instrumentation suggests minimal unsteadiness in the flow, less than a few percent of the mean (comparable to the sensor accuracy).

Wind tunnel testing, test techniques/methods, dyna↗

NASA NTRS1976

Review of coaxial flow gas core nuclear rocket fluid mechanics

The two areas where discussion and comparison of work is most useful are in (1) the factors which influence containment in cold flow studies; and (2) the effects of heat generation on containment. The work in these areas have not received any critical review in the past. The review is structured in such a way as to compare and contrast the related work of the program, rather than to preserve the chronological order of the work.

Weinstein, H.↗

DOE OSTI2025

Low-mode nonuniformity in direct-drive ICF implosions due to laser smoothing techniques employed on OMEGA

For successful laser-direct-drive inertial confinement fusion implosions, the laser irradiation must be highly uniform over the target surface. On OMEGA, multiple laser beams are used to illuminate targets quasi-uniformly. High-mode-number nonuniformities due to laser speckle on each individual beam are reduced by splitting each beam into two orthogonal polarizations (i.e., polarization smoothing, or PS) and a range of wavelengths (i.e., smoothing by spectral dispersion) that are dispersed at the target plane. However, cross-beam energy transfer (CBET) is sensitive to both the polarizations and wavelengths of the interacting beams, so the interplay between CBET and the laser-smoothing schemes results in unique intensity variation across each beam profile, which is a systematic source of low-mode drive nonuniformity on OMEGA. Here, we model these effects and find that the predicted ℓ = 1 mode in the laser-absorption distribution is consistent with the systematic core-flow direction that has been determined from the OMEGA implosion database. We also observe good agreement with the measured core-flow directions for two specific sets of implosions (one with PS, the other without PS) when we also account for the measured beam mispointing and the beam power imbalance.

Crossed beam scattering↗

NASA NTRS1991

Flow visualization studies of the internal flow characteristics in a simulated mixed flow vectored thrust ASTOVL engine configuration

The characteristics of internal flowfields in a Mixed Flow Vectored Thrust (MFVT) propulsion system for Short Take Off and Vertical Landing (STOVL) fighter aircraft were investigated using flow visualization techniques. A highly parametric 14 percent scale transparent model was used to simulate the MFVT propulsion system. The working medium was water. The mixing between the fan and core flow was studied using flow visualization for a variety of geometrical and flow parameters including three bypass ratio variations. The flow visualization technique involves the use of laser light sheet for illumination and fluorescent dye or air bubbles as tracers. Results indicate the existence of a strong vortex flowfield in the hover mode of operation that is responsible for good mixing between the fan and core flows.

Saripalli, K. R.↗

NASA NTRS1982

Experiments on high bypass internal mixer nozzle jet noise

Model scale jet noise data are presented for a variety of internal lobed mixer nozzle configurations for take off power settings in a static environment. The results are presented for a 17.5 cm diameter fan nozzle to show the effect on noise levels caused by changes in geometric shape of the internal, or core flow, nozzle. The geometric variables include the lobe discharge angle, the number of lobes, spacing between the center plug and lobe valley, lobe side wall shape and axial contour of the lobes. An annular plug core flow nozzle was also tested and is used as a baseline for comparative purposes. Comparison of data from the internal lobed configurations showed that for comparative effect, in terms of the effect on full scale perceived noise levels, was caused by a change in the lobe discharge angle. The results showed that increasing the discharge angle caused an increase as large as 7 dB in sound pressure levels in the high frequency portion of the spectra. Changes in the other geometric variables cause negligible effects.

Goodykoontz, J. H.↗

NASA NTRS1993

Small hydrogen/oxygen rocket flowfield behavior from heat flux measurements

The mixing and heat transfer phenomena in small rocket flow fields with fuel film cooling is not well understood. An instrumented, water-cooled chamber with a gaseous hydrogen/gaseous oxygen injector was used to gather steady-state inner and outer wall temperature profiles. The chamber was tested at 414 kPa (60 psia) chamber pressure, from mixture ratios of 3.41 to 8.36. Sixty percent of the fuel was used for film cooling. These temperature profiles were used as boundary conditions in a finite element analysis program, MSC/NASTRAN, to calculate the local radial and axial heat fluxes in the chamber wall. The normal heat fluxes were then calculated and used as a diagnostic of the rocket's flow field behavior. The normal heat fluxes determined were on the order of 1.0 to 3.0 MW/meters squared (0.6 to 1.8 Btu/sec-inches squared). In the cases where mixture ratio was 5 or above, there was a sharp local heat flux maximum in the barrel section of the chamber. This local maximum seems to indicate a reduction or breakdown of the fuel film cooling layer, possibly due to increased mixing in the shear layer between the film and core flows. However, the flow was thought to be completely laminar, as the throat Reynolds numbers were below 50,000 for all the cases. The increased mixing in the shear layer in the higher mixture ratio cases appeared not to be due to the transition of the flow from laminar to turbulent, but rather due to increased reactions between the hydrogen film and oxidizer-rich core flows.

Reed, Brian D.↗

NASA NTRS2001

Alleviation of Facility/Engine Interactions in an Open-Jet Scramjet Test Facility

Results of a series of shakedown tests to eliminate facility/engine interactions in an open-jet scramjet test facility are presented. The tests were conducted with the NASA DFX (Dual-Fuel eXperimental scramjet) engine in the NASA Langley Combustion Heated Scramjet Test Facility (CHSTF) in support of the Hyper-X program, The majority of the tests were conducted at a total enthalpy and pressure corresponding to Mach 5 flight at a dynamic pressure of 734 psf. The DFX is the largest engine ever tested in the CHSTF. Blockage, in terms of the projected engine area relative to the nozzle exit area, is 81% with the engine forebody leading edge aligned with the upper edge of the facility nozzle such that it ingests the nozzle boundary layer. The blockage increases to 95% with the engine forebody leading edge positioned 2 in. down in the core flow. Previous engines successfully tested in the CHSTF have had blockages of no more than 51%. Oil flow studies along with facility and engine pressure measurements were used to define flow behavior. These results guided modifications to existing aeroappliances and the design of new aeroappliances. These changes allowed fueled tests to be conducted without facility interaction effects in the data with the engine forebody leading edge positioned to ingest the facility nozzle boundary layer. Interaction effects were also reduced for tests with the engine forebody leading edge positioned 2 in. into the core flow, however some interaction effects were still evident in the engine data. A new shroud and diffuser have been designed with the goal of allowing fueled tests to be conducted with the engine forebody leading edge positioned in the core without facility interaction effects in the data. Evaluation tests of the new shroud and diffuser will be conducted once ongoing fueled engine tests have been completed.

Albertson, Cindy W.↗

NASA NTRS2014

Resonant Tidal Excitation of Internal Waves in the Earth's Fluid Core

It has long been speculated that there is a stably stratified layer below the core-mantle boundary, and two recent studies have improved the constraints on the parameters describing this stratification. Here we consider the dynamical implications of this layer using a simplified model. We first show that the stratification in this surface layer has sensitive control over the rate at which tidal energy is transferred to the core. We then show that when the stratification parameters from the recent studies are used in this model, a resonant configuration arrives whereby tidal forces perform elevated rates of work in exciting core flow. Specifically, the internal wave speed derived from the two independent studies (150 and 155 m/s) are in remarkable agreement with the speed (152 m/s) required for excitation of the primary normal mode of oscillation as calculated from full solutions of the Laplace Tidal Equations applied to a reduced-gravity idealized model representing the stratified layer. In evaluating this agreement it is noteworthy that the idealized model assumed may be regarded as the most reduced representation of the stratified dynamics of the layer, in that there are no non-essential dynamical terms in the governing equations assumed. While it is certainly possible that a more realistic treatment may require additional dynamical terms or coupling, it is also clear that this reduced representation includes no freedom for coercing the correlation described. This suggests that one must accept either (1) that tidal forces resonantly excite core flow and this is predicted by a simple model or (2) that either the independent estimates or the dynamical model does not accurately portray the core surface layer and there has simply been an unlikely coincidence between three estimates of a stratification parameter which would otherwise have a broad plausible range.

Geodynamo↗

DOE OSTI2020

Hybrid Ceramic-CMC Vane with EBC for Future Coal Derived Syngas Fired 65% Efficient Turbine Combined Cycle

The efficiency of both simple cycle and combined cycle power generation systems scale with the peak temperature at which the gas exits the combustor to drive the turbine. In conventional systems, a substantial fraction of the total turbine core flow exiting the compressor is diverted downstream to cool metallic turbine hardware rather than power the turbine, much of which is used to cool the first-stage turbine vane. The use of coal derived syngas fuels provides an additional challenge to the lifetime of materials utilized in the turbine, as particulate byproducts created in the coal gasification process melt in the combustion gas, and can subsequently deposit and interact with the turbine hardware. The development of durable hot-section materials capable of operating at temperature well above that of single crystal superalloy airfoil/zirconia based thermal barrier coatings is critical to realizing 65% efficient coal derived syngas fired gas turbine based power systems. To enable higher turbine inlet temperatures while lowering cooling air requirements, United Technologies Research Center (UTRC), the central R&D laboratory supporting UT Pratt & Whitney, led the conceptual design of a new type of ceramic composite turbine hot section materials system. The design focused on a novel hybrid monolithic ceramic-fiber reinforced ceramic matrix composite (CMC) first stage turbine vane having an environmental barrier coating. By utilizing ceramic construction in the turbine hot-section, the core flow normally used to cool metallic components will be substantially reduced, increasing efficiency and reducing emissions. To provide the framework for future demonstration testing, UTRC partnered with University of North Dakota Energy and Environmental Research Center (UNDEERC) to provide a conceptual design for a gasified coal fed high-pressure turbine combustor system designed to mimic the conditions expected in a future 65% fuel to busbar efficient syngas fueled gas turbine based combined cycle. The UNDEERC and UTRC collaborated on characterizing dusts from coal gasifier filtration systems.

10 SYNTHETIC FUELS↗