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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 325 records · Page 18
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Steady flow past sudden expansions at large Reynolds number. II - Navier-Stokes solutions for the cascade expansion

The equations of motion in the steady laminar flow past a sudden expansion at large Reynolds number, R, reduce to the boundary layer equations as R tends to the freestream valve, when the longitudinal length scale of the separated eddy increases linearly and indefinitely with R. A global Newton method is presently used to obtain finite difference solutions to the steady Navier-Stokes equations up to R of 1000 for a uniform inflow past a cascade of sudden expansions. For large expansion ratio values, the eddy length increases linearly with R; for smaller values, however, where boundary layer equation solutions could not be found, the steady solutions to the Navier-Stokes equations approach the limit of an inviscid eddy in length with increasing R.

Milos, Frank S.↗
NASA NTRS1989

Three-dimensional aerodynamics of an annular airfoil cascade including loading effects

A series of experiments are described which investigate and quantify the effect of loading on the three-dimensional flow through a subsonic annular cascade of cambered airfoils. At two levels of loading, detailed data quantify the cascade inlet velocity, the intrapassage flow field, the airfoil surface pressure distributions, the exit flow field, and the total pressure loss distributions. Aerodynamic loading is shown to strengthen the radial pressure gradient, the passage vortex structure, the vortex-endwall boundary layer interactions, and the losses.

Fleeter, S.↗
NASA NTRS1981

A rapid perturbation procedure for determining nonlinear flow solutions: Application to transonic turbomachinery flows

Perturbation procedures and associated computational codes for determining nonlinear flow solutions were developed to establish a method for minimizing computational requirements associated with parametric studies of transonic flows in turbomachines. The procedure that was developed and evaluated was found to be capable of determining highly accurate approximations to families of strongly nonlinear solutions which are either continuous or discontinuous, and which represent variations in some arbitrary parameter. Coordinate straining is employed to account for the movement of discontinuities and maxima of high gradient regions due to the perturbation. The development and results reported are for the single parameter perturbation problem. Flows past both isolated airfoils and compressor cascades involving a wide variety of flow and geometry parameter changes are reported. Attention is focused in particular on transonic flows which are strongly supercritical and exhibit large surface shock movement over the parametric range studied; and on subsonic flows which display large pressure variations in the stagnation and peak suction pressure regions. Comparisons with the corresponding 'exact' nonlinear solutions indicate a remarkable accuracy and range of validity of such a procedure.

Stahara, S. S.↗
NASA NTRS1983

Development of a multiple-parameter nonlinear perturbation procedure for transonic turbomachinery flows: Preliminary application to design/optimization problems

An investigation was conducted to continue the development of perturbation procedures and associated computational codes for rapidly determining approximations to nonlinear flow solutions, with the purpose of establishing a method for minimizing computational requirements associated with parametric design studies of transonic flows in turbomachines. The results reported here concern the extension of the previously developed successful method for single parameter perturbations to simultaneous multiple-parameter perturbations, and the preliminary application of the multiple-parameter procedure in combination with an optimization method to blade design/optimization problem. In order to provide as severe a test as possible of the method, attention is focused in particular on transonic flows which are highly supercritical. Flows past both isolated blades and compressor cascades, involving simultaneous changes in both flow and geometric parameters, are considered. Comparisons with the corresponding exact nonlinear solutions display remarkable accuracy and range of validity, in direct correspondence with previous results for single-parameter perturbations.

Stahara, S. S.↗
NASA NTRS1992

Lumley's energy cascade dissipation rate model for boundary-free turbulent shear flows

True dissipation occurs mainly at the highest wavenumbers where the eddy sizes are comparatively small. These high wavenumbers receive their energy through the spectral cascade of energy starting with the largest eddies spilling energy into the smaller eddies, passing through each wavenumber until it is dissipated at the microscopic scale. However, a small percentage of the energy does not spill continuously through the cascade but is instantly passed to the higher wavenumbers. Consequently, the smallest eddies receive a certain amount of energy almost immediately. As the spectral energy cascade continues, the highest wavenumber needs a certain time to receive all the energy which has been transferred from the largest eddies. As such, there is a time delay, of the order of tau, between the generation of energy by the largest eddies and the eventual dissipation of this energy. For equilibrium turbulence at high Reynolds numbers, there is a wide range where energy is neither produced by the large eddies nor dissipated by viscosity, but is conserved and passed from wavenumber to higher wavenumbers. The rate at which energy cascades from one wavenumber to another is proportional to the energy contained within that wavenumber. This rate is constant and has been used in the past as a dissipation rate of turbulent kinetic energy. However, this is true only in steady, equilibrium turbulence. Most dissipation models contend that the production of dissipation is proportional to the production of energy and that the destruction of dissipation is proportional to the destruction of energy. In essence, these models state that the change in the dissipation rate is proportional to the change in the kinetic energy. This assumption is obviously incorrect for the case where there is no production of turbulent energy, yet energy continues to cascade from large to small eddies. If the time lag between the onset on the energy cascade to the destruction of energy at the microscale can be modeled, then there will be a better representation of the dissipation process. Development of an energy cascade time scale equation is discussed.

Duncan, B. S.↗
NASA NTRS1999

Numerical Simulation of Complex Turbomachinery Flows

An unsteady, multiblock, Reynolds Averaged Navier Stokes solver based on Runge-Kutta scheme and Pseudo-time step for turbo-machinery applications was developed. The code was validated and assessed against analytical and experimental data. It was used to study a variety of physical mechanisms of unsteady, three-dimensional, turbulent, transitional, and cooling flows in compressors and turbines. Flow over a cylinder has been used to study effects of numerical aspects on accuracy of prediction of wake decay and transition, and to modify K-epsilon models. The following simulations have been performed: (a) Unsteady flow in a compressor cascade: Three low Reynolds number turbulence models have been assessed and data compared with Euler/boundary layer predictions. Major flow features associated with wake induced transition were predicted and studied; (b) Nozzle wake-rotor interaction in a turbine: Results compared to LDV data in design and off-design conditions, and cause and effect of unsteady flow in turbine rotors were analyzed; (c) Flow in the low-pressure turbine: Assessed capability of the code to predict transitional, attached and separated flows at a wide range of low Reynolds numbers and inlet freestream turbulence intensity. Several turbulence and transition models have been employed and comparisons made to experiments; (d) leading edge film cooling at compound angle: Comparisons were made with experiments, and the flow physics of the associated vortical structures were studied; and (e) Tip leakage flow in a turbine. The physics of the secondary flow in a rotor was studied and sources of loss identified.

Chernobrovkin, A. A.↗
NASA NTRS1974

Fluid mechanics, acoustics, and design of turbomachinery, part 1

A conference was conducted to analyze the effects of air flow on turbomachinery design. The subjects discussed are: (1) equations for compressible flow through turbomachines, (2) influence of axial velocity ratio on cascade performance, (3) three dimensional flow in transonic axial compressor blade rows, (4) prediction of turbulent shear layers in turbomachines, and (5) boundary layers in centrifugal compressors.

Lakshminarayana, B.↗
NASA NTRS1981

Sound generated in a cascade by three-dimensional disturbances convected in a subsonic flow

Discrete tone sound generation in a subsonic fan subject to three-dimensional disturbances is investigated. The analytical model used treats the fan rotor and stator as linear cascades of thin airfoils in a rectangular duct subject to a three-dimensional gust for which a complete aerodynamic theory already exists. The sound pressure can then be cast as the sum of a finite number of discrete sound waves (modes) the magnitude of which depends on an unknown function satisfying a singular integral equation. Similarity rules are derived to reduce the problem to that of a two-dimensional gust. Three-dimensional effects on the cut-off condition, the sound pressure, and the acoustic power are first investigated for each mode. The theory is then applied to noise generated by typical rotor-wake-defect and rotor-tip-vortex disturbances interacting with a stator.

Atassi, H.↗
NASA NTRS1975

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

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.

Dietrich, D. A.↗
NASA NTRS1972

Thermochemical nonequilibrium in atomic hydrogen at elevated temperatures

A numerical study of the nonequilibrium flow of atomic hydrogen in a cascade arc was performed to obtain insight into the physics of the hydrogen cascade arc. A rigorous mathematical model of the flow problem was formulated, incorporating the important nonequilibrium transport phenomena and atomic processes which occur in atomic hydrogen. Realistic boundary conditions, including consideration of the wall electrostatic sheath phenomenon, were included in the model. The governing equations of the asymptotic region of the cascade arc were obtained by writing conservation of mass and energy equations for the electron subgas, an energy conservation equation for heavy particles and an equation of state. Finite-difference operators for variable grid spacing were applied to the governing equations and the resulting system of strongly coupled, stiff equations were solved numerically by the Newton-Raphson method.

Scott, R. K.↗
NASA NTRS1993

Computation of flows in a turn-around duct and a turbine cascade using advanced turbulence models

Numerical investigation has been carried out to evaluate the capability of the Algebraic Reynolds Stress Model (ARSM) and the Nonlinear Stress Model (NLSM) to predict strongly curved turbulent flow in a turn-around duct (TAD). The ARSM includes the near-wall damping term of pressure-strain correlation phi(sub ij,w), which enables accurate prediction of individual Reynolds stress components in wall flows. The TAD mean flow quantities are reasonably well predicted by various turbulence models. The ARSM yields better predictions for both the mean flow and the turbulence quantities than the NLSM and the k-epsilon (k = turbulent kinetic energy, epsilon = dissipation rate of k) model. The NLSM also shows slight improvement over the k-epsilon model. However, all the models fail to capture the recovery of the flow from strong curvature effects. The formulation for phi(sub ij,w) appears to be incorrect near the concave surface. The hybrid k-epsilon/ARSM, Chien's k-epsilon model, and Coakley's q-omega (q = the square root of k, omega = epsilon/k) model have also been employed to compute the aerodynamics and heat transfer of a transonic turbine cascade. The surface pressure distributions and the wake profiles are predicted well by all the models. The k-epsilon model and the k-epsilon/ARSM model provide better predictions of heat transfer than the q-omega model. The k-epsilon/ARSM solutions show significant differences in the predicted skin friction coefficients, heat transfer rates and the cascade performance parameters, as compared to the k-epsilon model. The k-epsilon/ARSM model appears to capture, qualitatively, the anisotropy associated with by-pass transition.

Lakshminarayana, B.↗
NASA NTRS1950

The Development of a Hollow Blade for Exhaust Gas Turbines

The subject of the development of German hollow turbine blades for use with internal cooling is discussed in detail. The development of a suitable blade profile from cascade theory is described. Also a discussion of the temperature distribution and stresses in a turbine blade is presented. Various methods of manufacturing hollow blades and the methods by which they are mounted in the turbine rotor are presented in detail.

TURBINES - BLADES, HOLLOW↗
NASA NTRS1957

Aerodynamic Forces on a Vibrating Unstaggered Cascade

The unsteady aerodynamic forces, [based on two-dimensional incompressible flow considerations], are determined for an unstaggered cascade, the blades of which are vibrating in phase in an approach flow parallel to the blades.

Soehngen, H.↗
NASA NTRS1977

A two-dimensional cascade solution using minimized surface singularity density distributions - with application to film cooled turbine blades

An investigation of the effects of coolant injection on the aerodynamic performance of cooled turbine blades is presented. The coolant injection is modeled in the inviscid irrotational adiabatic flow analysis through the cascade using the distributed singularities approach. The resulting integral equations are solved using a minimized surface singularity density criteria. The aerodynamic performance was evaluated using this solution in conjunction with an existing mixing theory analysis. The results of the present analysis are compared with experimental measurements in cold flow tests.

Mcfarland, E.↗
NASA NTRS1996

Experimental investigation of unsteady flows at large incidence angles in a linear oscillating cascade

The aerodynamics of a cascade of airfoils oscillating in torsion about the midchord is investigated experimentally at a large mean incidence angle and, for reference, at a low mean incidence angle. The airfoil section is representative of a modern, low aspect ratio, fan blade tip section. Time-dependent airfoil surface pressure measurements were made for reduced frequencies up to 0.8 for out-of-phase oscillations at Mach numbers up to 0.8 and chordal incidence angles of 0 deg and 10 deg. For the 10 deg chordal incidence angle, a separation bubble formed at the leading edge of the suction surface. The separated flow field was found to have a dramatic effect on the chordwise distribution of the unsteady pressure. In this region, substantial deviations from the attached flow data were found with the deviations becoming less apparent in the aft region of the airfoil for all reduced frequencies. In particular, near the leading edge the separated flow had a strong destabilizing influence while the attached flow had a strong stabilizing influence.

Buffum, Daniel H.↗
NASA NTRS1989

The measurement of boundary layers on a compressor blade in cascade. Volume 1: Experimental technique, analysis and results

Measurements were made of the boundary layers and wakes about a highly loaded, double-circular-arc compressor blade in cascade. These laser Doppler velocimetry measurements have yielded a very detailed and precise data base with which to test the application of viscous computational codes to turbomachinery. In order to test the computational codes at off-design conditions, the data were acquired at a chord Reynolds number of 500,000 and at three incidence angles. Moreover, these measurements have supplied some physical insight into these very complex flows. Although some natural transition is evident, laminar boundary layers usually detach and subsequently reattach as either fully or intermittently turbulent boundary layers. These transitional separation bubbles play an important role in the development of most of the boundary layers and wakes measured in this cascade and the modeling or computing of these bubbles should prove to be the key aspect in computing the entire cascade flow field. In addition, the nonequilibrium turbulent boundary layers on these highly loaded blades always have some region of separation near the trailing edge of the suction surface. These separated flows, as well as the subsequent near wakes, show no similarity and should prove to be a challenging test for the viscous computational codes.

Zierke, William C.↗
NASA NTRS1981

Semi-actuator disk theory for compressor choke flutter

A mathematical anaysis predict the unsteady aerodynamic utilizing semi actuator theory environment for a cascade of airfoils harmonically oscillating in choked flow was developed. A normal shock is located in the blade passage, its position depending on the time dependent geometry, and pressure perturbations of the system. In addition to shock dynamics, the model includes the effect of compressibility, interblade phase lag, and an unsteady flow field upstream and downstream of the cascade. Calculated unsteady aerodynamics were compared with isolated airfoil wind tunnel data, and choke flutter onset boundaries were compared with data from testing of an F100 high pressure compressor stage.

Micklow, J.↗