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Results for “Conditional Normalizing Flow”

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At least 325 records · Page 18

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NASA NTRS1978

Wind tunnel tests of a blade subjected to midchord torsional oscillation at high subsonic stall flutter conditions

A mechanical drive system for oscillating blades in a wind tunnel at frequencies up to 767 hertz and amplitudes of + or - 1.2 deg is described. High-speed motion pictures of schlieren images of the flow over a double-circular arc blade oscillating in harmonic motion about the midchord revealed extensive shock patterns at a nominal free stream Mach number of 0.7, a mean angle of attack of 4 deg, and reduced frequency of about 0.7. A phase lag resulting from the slow response of the flow to the motion of the blade increased with increasing reduced frequency. This phase lag, based on the difference between the time the blade attained its maximum angle of attack and the time required for the normal shock to reach its extreme downstream position, was nominally 100 deg at the above conditions.

Boldman, D. R.↗

NASA NTRS1975

Critical flow and pressure ratio data for LOX flowing through nozzles

LOX and LN2 data for two-phase critical flow through nozzles were acquired with precision control. The principal measured parameters were inlet conditions, critical flow rate and critical flow pressure ratio. It is conclusively demonstrated that the principle of corresponding states can be applied to two-phase choked flow through nozzles. It is also shown that the proper normalizing parameters are developed, and that current theories can provide an adequate means for extrapolating data to other fluids.

Hendricks, R. C.↗

NASA NTRS1975

Critical flow and pressure ratio data for LOX flowing through nozzles

LOX and LN2 data for two-phase critical flow through nozzles have been acquired with precision control. The principal measured parameters were inlet conditions, critical flow rate and critical flow pressure ratio. The data conclusively demonstrate that the principle of corresponding states can be applied to two-phase choked flow through nozzles. These data also demonstrate that the proper normalizing parameters have been developed and current theories can provide an adequate means for extrapolating to other fluids.

Hendricks, R. C.↗

NASA NTRS1981

Multivalued velocity field model of turbulence

The considered study is based on new theoretical concepts regarding a post-instability model of a fluid discussed by Zak (1980). The model permits the completion of the governing equations of turbulence by introducing multivalued fields of velocities. Attention is given to the mechanism of energy dissipation, the characteristic wave propagation, a simplified model, the formation of turbulence around stagnation points, the formulation of boundary conditions, and the mechanism of turbulence formation. The mechanism of turbulence formation can be understood as propagation of initial discontinuities from the boundaries into a flow with the characteristic velocity which is defined by the normal (to the boundary) velocity components. These components emerge at the boundary as a result of jumps in the tangential components due to the continuity equation.

Zak, M.↗

NASA NTRS1981

Interaction of upstream flow distortions with high Mach number cascades

Features of the interaction of flow distortions, such as gusts and wakes with blade rows of advance type fans and compressors having high tip Mach numbers are modeled. A typical disturbance was assumed to have harmonic time dependence and was described, at a far upstream location, in three orthogonal spatial coordinates by a double Fourier series. It was convected at supersonic relative to a linear cascade described as an unrolled annulus. Conditions were selected so that the component of this velocity parallel to the axis of the turbomachine was subsonic, permitting interaction between blades through the upstream as well as downstream flow media. A strong, nearly normal shock was considered in the blade passages which was allowed curvature and displacement. The flows before and after the shock were linearized relative to uniform mean velocities in their respective regions. Solution of the descriptive equations was by adaption of the Wiener-Hopf technique, enabling a determination of distortion patterns through and downstream of the cascade as well as pressure distributions on the blade and surfaces. Details of interaction of the disturbance with the in-passage shock were discussed. Infuences of amplitude, wave length, and phase of the disturbance on lifts and moments of cascade configurations are presented. Numerical results are clarified by reference to an especially orderly pattern of upstream vertical motion in relation to the cascade parameters.

Englert, G. W.↗

NASA NTRS1982

Dynamic load measurements with delta wings undergoing self-induced roll-oscillations

The aerodynamic forces acting on a delta wing, mounted on a free-to-roll sting-balance apparatus, were measured. Two wing planforms having leading edge sweeps of 76 and 80 deg were tested, but only the wing with the 80 deg sweep would undergo periodic self-induced roll oscillation. The time dependent forces and roll angles for this wing were then recorded for various test conditions. In these tests a considerable drop in the average normal force of the free-to-roll wing was measured, relative to the normal force obtained in the static tests. Also, the helium-bubble flow visualization technique was used to gain some insight into the periodic motion of the separated leading edge vortices.

Levin, D.↗

NASA NTRS1984

SSME thrust chamber simulation using Navier-Stokes equations

The capability of the PHOENICS fluid dynamics code in predicting two-dimensional, compressible, and reacting flow in the combustion chamber and nozzle of the space shuttle main engine (SSME) was evaluated. A non-orthogonal body fitted coordinate system was used to represent the nozzle geometry. The Navier-Stokes equations were solved for the entire nozzle with a turbulence model. The wall boundary conditions were calculated based on the wall functions which account for pressure gradients. Results of the demonstration test case reveal all expected features of the transonic nozzle flows. Of particular interest are the locations of normal and barrel shocks, and regions of highest temperature gradients. Calculated performance (global) parameters such as thrust chamber flow rate, thrust, and specific impulse are also in good agreement with available data.

Przekwas, A. J.↗

NASA NTRS1987

Remote noncontacting measurements of heat transfer coefficients for detection of boundary layer transition in wind tunnel tests

An infrared measurement system is used that consists of a laser heating source, an infrared camera for data acquisition, and a video recorder for data storage. A laser beam is scanned over an airfoil, heating its surface to a few degrees above ambient. An infrared camera then measures the temperature of the airfoil over a two-dimensional field, and these temperatures are stored as a function of time on a video recorder. The resulting temperature pictures are digitized and an iterative approximation algorithm is used to extract the heat transfer coefficient. The resulting values are normalized to the natural convection condition. The technique has been applied in low-speed wind tunnel tests and compared to well-established hot-film measurements which were made simultaneously to confirm the flow conditions. Heat transfer coefficients were determined using a linear scanning pattern, to indicate the position of natural and of artificially induced transition on an airfoil, at various wind speeds. The technique is shown to be sensitive to transition at low Mach numbers. The advantages of the technique are discussed.

Heath, D. Michele↗

NASA NTRS1946

Investigation of Operating Characteristics of an Engine Equipped with Modifications to Eliminate Fuel-Evaporation Icing

Two modified fuel-injection systems, a drilled-inducer type and a spinner type, that prevent serious fuel-evaporation icing were installed on a V-type, liquid-cooled aircraft engine and a preliminary investigation was conducted to determine the effect on engine operating characteristics. The spinner system was also ground - and flight tested on a twin-engine fighter airplane. Flight measurements of cylinder-head temperature over a range of fuel-air ratios and engine power conditions were made at an altitude of approximately 10,000 feet. Starting and accelerating of the engine on the ground were unaffected by the fuel-injection modifications. During the flight investigation, no appreciable variation occurred between the maximum and minimum cylinder-head temperature with the standard and modified system for the same power condition and no irregularity of mixture distribution could be detected throughout the power range of the engine. Normal mixture distribution was also indicated by a similar response of cylinder-head temperature for variations of fuel-air ratio at manifold pressures of 25 and 35 inches of mercury absolute. Both modified fuel-injection systems required less fuel-nozzle pressure than the standard system to obtain the desired fuel-air ratio for given air-flow condition.

Mulholland, Donald R.↗

NASA NTRS2007

Implementation of In-Situ Impedance Techniques on a Full Scale Aero-Engine System

Determination of acoustic liner impedance for jet engine applications remains a challenge for the designer. Although suitable models have been developed that take account of source amplitude and the local flow environment experienced by the liner, experimental validation of these models has been difficult. This is primarily due to the inability of researchers to faithfully mimic the environment in jet engine nacelles in the laboratory. An in-situ measurement technique, one that can be implemented in an actual engine, is desirable so an accurate impedance can be determined for future modeling and quality control. This paper documents the implementation of such a local acoustic impedance measurement technique that is used under controlled laboratory conditions as well as on full scale turbine engine liner test article. The objective for these series of in-situ measurements is to substantiate treatment design, provide understanding of flow effects on installed liner performance, and provide modeling input for fan noise propagation computations. A series of acoustic liner evaluation tests are performed that includes normal incidence tube, grazing incidence tube, and finally testing on a full scale engine on a static test stand. Lab tests were intended to provide insight and guidance for accurately measuring the impedance of the liner housed in the inlet of a Honeywell Tech7000 turbofan. Results have shown that one can acquire very reasonable liner impedance data for a full scale engine under realistic test conditions. Furthermore, higher fidelity results can be obtained by using a three-microphone coherence technique that can enhance signal-to-noise ratio at high engine power settings. This research has also confirmed the limitations of this particular type of in-situ measurement. This is most evident in the installation of instrumentation and its effect on what is being measured.

Gaeta, R. J.↗

NASA NTRS1991

One-way nesting for a primitive equation ocean model

Prognostic numerical models for atmospheric and oceanic circulations require initial fields, boundary conditions, and forcing functions in addition to a consistent set of partial differential equations, including a state relation and equations expressing conservation of mass, momentum, and energy. Depending on the horizontal domain to be modeled, the horizontal boundary conditions are either physically obvious or extremely difficult to specify consistently. If the entire atmosphere is modeled, periodic horizontal boundary conditions are appropriate. On the other hand, the physical horizontal boundaries on the entire ocean are solid walls. Obviously, the normal velocity at a solid wall is zero while the specification of the tangential velocity depends on the mathematical treatment of the horizontal viscous terms. Limitations imposed by computer capacity and cost, as well as research interests, have led to the use of limited area models to study flows in the atmosphere and ocean. The limited area models do not have physical horizontal boundaries, merely numerical ones. Correctly determining these open boundary conditions for limited-area numerical models has both intrigued and frustrated numerical modelers for decades. One common approach is to use the closed or solid wall boundary conditions for a limited-area model. The argument given for this approach is that the boundary conditions affect flow near the walls but that none of these effects are propagated into the interior. Therefore, one chooses a big enough domain that the central region of interest is not corrupted by the boundary flow. Research in progress to model the North Atlantic circulation vividly illustrates the pitfalls of this approach. Two model runs are compared: (1) the southern boundary at 20S between latitudes 0 and 40W is artificially closed; and (2) the same boundary is specified as open with an inward transport of 15 Sv (determined from a global model with the same physics) uniformly spread across the boundary. A comparison of both runs is presented.

Blake, D. W.↗

NASA NTRS1992

In-flight leading-edge extension vortex flow-field survey measurements on a F-18 aircraft at high angle of attack

Flow-field measurements on the leading-edge extension (LEX) of the F-18 High Alpha Research Vehicle (HARV) were obtained using a rotating rake with 16 hemispherical-tipped five-hole probes. Detailed pressure, velocity, and flow direction data were obtained through the LEX vortex core. Data were gathered during 1-g quasi-stabilized flight conditions at angles of attack alpha from 10 degrees to 52 degrees and at Reynolds numbers based on mean aerodynamic cord up to 16 x 10(exp 6). Normalized dynamic pressures and crossflow velocities clearly showed the primary vortex above the LEX and formation of a secondary vortex at higher angles of attack. The vortex was characterized by a ring of high dynamic pressure surrounding a region of low dynamic pressure at the vortex core center. The vortex core, subcore diameter, and vertical location of the core above the LEX increased with angle of attack. Minimum values for static pressure were obtained in the vortex subcore and decreased nearly linearly with increasing angle of attack until vortex breakdown. Rake-measured static pressures were consistent with previously documented surface pressures and showed good agreement with flow visualization flight test results. Comparison of the LEX vortex flight test data to computational solutions at alpha approximately equals 19 degrees and 30 degrees showed fair correlation.

David M Richwine↗

NASA NTRS2000

Propulsion Aeroelastic Analysis Developed for Flutter and Forced Response

The NASA Glenn Research Center at Lewis Field develops new technologies to increase the fuel efficiency of aircraft engines, improve the safety of engine operation, reduce emissions, and reduce engine noise. With the development of new designs for fans, compressors, and turbines to achieve these goals, the basic aeroelastic requirements are that there should be no flutter (self-excited vibrations) or high resonant blade stresses (due to forced response) in the operating regime. Therefore, an accurate prediction and analysis capability is required to verify the aeroelastic soundness of the designs. Such a three-dimensional viscous propulsion aeroelastic analysis capability has been developed at Glenn with support from the Advanced Subsonic Technology (AST) program. This newly developed aeroelastic analysis capability is based on TURBO, a threedimensional unsteady aerodynamic Reynolds-averaged Navier-Stokes turbomachinery code developed previously under a grant from Glenn. TURBO can model the viscous flow effects that play an important role in certain aeroelastic problems such as flutter with flow separation, flutter at high loading conditions near the stall line (stall flutter), flutter in the presence of shock and boundary-layer interaction, and forced response due to wakes and shock impingement. In aeroelastic analysis, the structural dynamics representation of the blades is based on normal modes. A finite-element analysis code is used to calculate these in-vacuum vibration modes and the associated natural frequencies. In an aeroelastic analysis using the TURBO code, flutter and forced response are modeled as being uncoupled. To calculate if a blade row will flutter, one prescribes the motion of the blade to be a harmonic vibration in a specified in-vacuum normal mode. An aeroelastic analysis preprocessor is used to generate the displacement field required for the analysis. The work done by aerodynamic forces on the vibrating blade during a cycle of vibration is calculated. If this work is positive, the blade is dynamically unstable, since it will extract energy from the flow, leading to an increase in the blade s oscillation amplitude. The forced-response excitations on a blade row are calculated by modeling the flow through two adjacent blade rows using the TURBO code. The blades are assumed to be rigid. As an option, a single blade row can be modeled with the upstream blade row influence represented by a time-varying disturbance (gust) at the inlet boundary. The unsteady forces on a blade row from such analyses are used in a structural analysis along with the blade structural dynamics characteristics and aerodynamic damping associated with blade vibration to calculate the resulting dynamic stresses on the blade.

Bakhle, Milind A.↗

NASA NTRS1981

Measurements in ducted flows by laser-Doppler anemometry

The merits of small-scale, water-flow experiments with measurements by laser-Doppler anemometry are indicated in relation to previously reported results. New measurements are reported in a square-to-round transition duct and show that, for example, the pressure forces overwhelm any small normal-stress driven secondary flows in the inlet plane to given maximum values of around 0.07 of the bulk velocity in the exit plane. The precision of manufacture of the ducts means that the geometric boundary conditions are accurately known and the results are therefore suitable for calculation purposes. Similarly, the instrumentation allows the flow boundary conditions to be accurately determined.

Taylor, A. M. K. P.↗

NASA NTRS1986

Magnetized thermal conduction fronts

The evolution of planar thermal conduction fronts in the presence of a dynamically weak, but otherwise self-consistent, magnetic field is considered. The field is assumed to be connected and untangled. In the diffusion limit for the thermal conductivity, these fronts exhibit self-similar behavior, even in the presence of a field. The role of the field is restricted to channeling the heat flux along its lines of force, and it enters into the problem as a dimensionless angle variable. 'Combing' (or opening) of insulating field lines by the evaporative flow is explicitly demonstrated. Unless the field is nearly perpendicular to the front normal in the hot gas, insulating effects are not profound. Self-similarity breaks down if the front becomes saturated, and under certain conditions magnetized saturated conduction fronts cannot propagate: the solution characteristics of the wave equation form caustics. The physical resolution is the advent of two-fluid (nonlocal) heating. Such Coulomb-heated fronts are expected to be relatively rare in typical astrophysical systems. The large-scale effects of a magnetic field on cloud evaporation in the interstellar medium are briefly discussed, and it is suggested that these fields preclude the presence of time-independent evaporative solutions. Thermal interfaces may then continue to evolve until radiative cooling halts their development; large tracts of warm 10,000 K gas may result.

Balbus, S. A.↗

NASA NTRS1992

Kinetic structure of slow shocks - Effects of the electromagnetic ion/ion cyclotron instability

The structure of slow magnetosonic shocks in the low beta regime is analyzed with attention given to ion heating and the effects of waves upstream of the electromagnetic ion/ion cyclotron (EMIIC) instability. Shock formation is assessed by means of three methods - a relaxation method and two based on dynamic flow interactions - to determine the effects of initialization and boundary conditions on the formation. Good solutions are found with the piston method and the similar flow-flow method in which the plasma is injected from two boundaries to form two slow shocks. Plasma parameters and shock normal angle are found to be the key variables dictating the structure of the magnetosonic shocks. Four unique classes of resultant shock structures are described in which classical, steady, or nonsteady behavior is found. The analysis also yields insight into the relationship between EMIIC instability and ion dissipation.

Omidi, N.↗

NASA NTRS1996

Fluid Creep Effects on Near-Wall Solute Transport for Non-Isothermal Ampoules

There is a growing practical and theoretical interest in developing accurate macroscopic modelling for flows arising in chemical or physical vapor transport (VT) crystal growth experiments, including those conducted in reduced gravity environments. Rosner was the first person to point out that previously neglected rarefield gas dynamics phenomena (Stefan and bouyancy-driven flows) become rather important sources of convection. In particular, the combination of rarefaction and strong gradients of temperature (and/or concentration) tangential to the side-walls of the ampoule induces convective flows known as thermal (and concentration) 'creep' respectively. His order-of-magnitude estimates revealed that thermal creep effects can be non-negligible even at normal gravitational levels. On the macroscopic level, the bulk fluid mechanics can be adequately described by the familiar macroscopic equations as long as the boundary conditions are modified to account for the integrated effect of kinetic boundary layers adjacent to solid boundaries. Motivated by the growing importance of these phenomena, we have embarked on a series of computational studies to elucidate these fundamental creep-induced effects for a rarefied gas in simple, two-dimensional confined geometries. However, unlike previous related studies, we resort to a microscopic description of the gas, mathematically expressed by the Boltzmann integro-differential equation. We employ the direct simulation Monte Carlo (DSMC) method of Bird, the theoretical foundations and several practical applications. In the case of thermally induced flows, the no-time counter method of Bird is used, as implemented for a hard-sphere gas. The scheme has been also extended to account for realistic molecular interaction models, an extension necessary if the diffusion physics underlying concentration creep are to be captured.

Papadopoulos, Dimitrios↗

NASA NTRS1973

Unsteady subsonic compressible flow around finite thickness wings.

A general formulation for the unsteady subsonic compressible potential flow around aircraft having arbitrary configurations is presented. An integral representation of the velocity potential is obtained. From this a linear integral equation relating the perturbation potential and its normal derivative (which is known from the boundary conditions) is derived. For the numerical solution of the integral equation, the surface of the aircraft is divided into small elements and the potential is assumed to be constant within each element. Numerical results are obtained for an oscillating finite-thickness wing and indicate good convergence and excellent agreement with existing lifting surface solutions.

Morino, L.↗