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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 379 records · Page 21
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Forcing function modeling for flow induced vibration

The fundamental forcing function unsteady aerodynamics for application to turbomachine blade row forced response are considered, accomplished through a series of experiments performed in a rotating annular cascade and a research axial flow turbine. In particular, the unsteady periodic flowfields downstream of rotating rows of perforated plates, airfoils and turbine blade rows are measured with a cross hot-wire and an unsteady total pressure probe. The unsteady velocity and static pressure fields were then analyzed harmonically and split into vortical and potential gusts, accomplished by developing a gust splitting analysis which includes both gust unsteady static pressure and velocity data. The perforated plate gusts closely were found to be linear theory vortical gusts, satisfying the vortical gust constraints. The airfoil and turbine blade row generated velocity perturbations did not satisfy the vortical gust constraints. However, the decomposition of the unsteady flow field separated the data into a propagating vortical component which satisfied these vortical gust constraints and a decaying potential component.

Fleeter, Sanford↗
NASA NTRS1974

Laser Doppler velocimeter measurements in a turbine stator cascade facility

A laser Doppler velocimeter (LDV) developed for mapping the flow velocity downstream from a 32-inch diameter annular cascade of turbine stator vanes in described. The LDV measurements were taken in a plane located approximately 0.5 inch downstream of the trailing edges of the vanes. Two components of the mean velocity (axial and circumferential) were measured. The flow velocities were in the high subsonic range. The LDV optics are of the dual scatter type with off-axis collection of the scattered light. The electronics system is based on the measurement of the time interval corresponding to eight periods of the Doppler signal and has a range of 10 to 80 MHz. The LDV measurements are compared with previous measurements made with a pressure probe.

Seasholtz, R. G.↗
NASA NTRS1974

Laser Doppler velocimeter measurements in a turbine stator cascade facility

A laser Doppler velocimeter (LDV) developed for mapping the flow velocity downstream from a 32-inch diameter annular cascade of turbine stator vanes is described. The LDV measurements were taken in a plane located approximately 0.5 inch downstream of the trailing edges of the vanes. Two components of the mean velocity (axial and circumferential) were measured. The flow velocities were in the high subsonic range. The LDV optics are of the dual scatter type with off-axis collection of the scattered light. The electronics system is based on the measurement of the time interval corresponding to eight periods of the Doppler signal and has a range of 10 to 80 MHz. The LDV measurements are compared with previous measurements made with a pressure probe.

Seasholtz, R. G.↗
NASA NTRS2007

Boundary Layer Development on a Turbine Blade in a Linear Cascade

Several different boundary-layer development patterns for flow over the suction surface of a turbine airfoil in a linear cascade were studied and documented using a sliding surface hot-film sensor. The state of the boundary layer, whether laminar, transitional or turbulent, was determined at numerous locations along the airfoil suction surface from leading to trailing edge. Boundary-layer transition from laminar to turbulent flow through laminar separation and turbulent reattachment, or through a combination of bypass transition and strong and weak separation and turbulent reattachment, or through solely bypass transition without separation, was observed and benchmark data were recorded. Surface flow visualization and numerical boundary-layer analysis results are consistent with the hot-film data. Flow and geometry information necessary for nmerical code operation is available.

Halstead, Dave↗
NASA NTRS1986

Chaotic motion in the Jovian atmosphere

Strong nonlinear interactions among unstable waves and the mean flow occur in a simplified quasigeostrophic spectral model of the upper troposphere of Jupiter. The upper boundary of the layer inhibits vertical motion while at the lower boundary perturbations of the potential temperature are not permitted. On an infinite beta plane the forced flow of alternating zones of prograde and retrograde zonal winds, decreasing with height, are linearly unstable and it is shown that the nonlinear terms stabilize the flow by bounding the growth of the eddies. Explicit viscosity terms are not needed. This does not imply that energy would not cascade to the small scale flow but suggests that the nature of the large scale flow is independent of the viscosity at small scales. Numerical time integration shows the flow to be chaotic but, in some cases, with transient propagating features and meandering zonal flow.

Pirraglia, Joseph↗
NASA NTRS1980

Experimental determination of unsteady blade element aerodynamics in cascades. Volume 1: Torsion mode cascade

A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic torsional flutter. This five bladed cascade had a solidity of 1.17 and a setting angle of 1.07 rad. Graphite epoxy airfoils were fabricated to achieve the realistically high reduced frequency level of 0.44. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time-steady and time-unsteady flow field surrounding the center cascade airfoil were investigated. The effects of reduced solidity and decreased setting angle on the flow field were also evaluated.

Riffel, R. E.↗
NASA NTRS1991

Wind tunnel wall effects in a linear oscillating cascade

Experiments in a linear oscillating cascade reveal that the wind tunnel walls enclosing the airfoils have, in some cases, a detrimental effect on the oscillating cascade aerodynamics. In a subsonic flow field, biconvex airfoils are driven simultaneously in harmonic, torsion-mode oscillations for a range of interblade phase angle values. It is found that the cascade dynamic periodicity - the airfoil to airfoil variation in unsteady surface pressure - is good for some values of interblade phase angle but poor for others. Correlation of the unsteady pressure data with oscillating flat plate cascade predictions is generally good for conditions where the periodicity is good and poor where the periodicity is poor. Calculations based upon linearized unsteady aerodynamic theory indicate that pressure waves reflected from the wind tunnel walls are responsible for the cases where there is poor periodicity and poor correlation with the predictions.

Buffum, Daniel H.↗
NASA NTRS1992

Flutter analysis of a supersonic cascade in time domain using an ADI Euler solver

The aeroelastic stability of a two-dimensional cascade oscillating in supersonic axial flow is analyzed in the time domain. The aeroelastic model consists of a single degree of freedom typical section structural model for each blade of the cascade and an unsteady two-dimensional cascade aerodynamic model based on the Euler equations. The Euler equations are solved using a time accurate Alternating Direction Implicit (ADI) solution scheme. The aeroelastic equations are integrated in time. The effect of interblade phase angle is included in the aeroelastic analysis by an appropriate choice of initial and boundary conditions. Flutter predictions are obtained from the time response of a flat plate cascade in single degree of freedom pitching motion. The results correlate well with those obtained from a separate frequency domain flutter analysis for all values of interblade phase angles considered. Flutter results are then presented for cascades having airfoil sections representative of a supersonic throughflow fan. The validity of the time integration method for a cascade of airfoils at various interblade phase angles is demonstrated.

Reddy, T. S. R.↗
NASA NTRS1982

An experience in mesh generation for three-dimensional calculation of potential flow around a rotating propeller

An attempt is made to develop a three-dimensional, finite volume computational code for highly swept, twisted, small aspect ratio propeller blades with supersonic tip speeds, in a way that accounts for cascade effects, hub-induced flow, and nonlinear transonic effects. Attention is presently given to the generation of a computational mesh for such a complex propeller configuration, with the aim of sharing developmental process experience. The problem treated is unique, in that blade chord, blade length, hub length and blade-to-blade distance represent several characteristic length scales among which there is considerable disparity. An ad hoc mesh-generation scheme is accordingly developed.

Jou, W.-H.↗
NASA NTRS1972

Systematic two-dimensional cascade tests. Volume 3: Slotted double circular-arc hydrofoils

Performance parameters are presented for cascades of slotted double circular-arc hydrofoils tested over a range of systematically introduced variables in a rectilinear cascade tunnel which uses water as the test medium. Cascade configurations included various combinations of an inlet flow angle of 50, 60 and 70 deg; a cascade solidity of 0.75, 1.00 and 1.50; a hydrofoil camber angle of 20, 30, 40 and 45 deg; and angles of incidence between positive and negative stall. The slot was positioned at the 45 percent chord station and the slot exit width was 0.047-in. Tests were also performed with the slot positioned at the 35 percent chord station and with slot widths of 0.63 and 0.094-in. These data were correlated to indicate the effects of slot location and slot width on minimum loss incidence and deviation angles. In addition, a comparison is presented of the performance parameters for cascades of slotted and unslotted hydrofoils.

Columbo, R. M.↗
NASA NTRS1974

Numerical solution of the time-dependent compressible Navier-Stokes equations in inlet regions

The results of a study to determine the effects of compressibility on the viscous flow through channels that have straight, parallel walls are presented. Two channel configurations are considered, the flow between two semi-infinite flat plates with uniform flow prescribed at the inlet plane and a cascade of semi-infinite flat plates with uniform flow introduced upstream. The flow field is modeled by using the time dependent, compressible Navier-Stokes equations. Time dependent solutions are obtained by using an explicit finite difference technique which advances the pressure on near field subsonic boundaries such that accurate steady state solutions are obtained. Steady state results at Reynolds number 20 and 150 are presented for Mach numbers between 0.09 and 0.36 and compared with the incompressible solutions of previous studies.

Olson, L. E.↗
NASA NTRS1957

Comparison of NACA 65-series compressor-blade pressure distributions and performance in a rotor and in cascade

An investigation has been conducted to compare the performance of NACA 65-series compressor blades in two-dimensional cascade with that in an axial flow compressor. Blade pressure distributions were obtained by the use of a mercury-seal pressure-transfer device. The comparison indicated that cascade data accurately predicted the turning angle and blade pressure distribution obtained in the compressor at design conditions.

CASCADES↗
NASA NTRS1981

The experimental verification of a streamline curvature numerical analysis method applied to the flow through an axial flow fan

The experimental verification of an inviscid, incompressible through-flow analysis method is presented. The primary component of this method is an axisymmetric streamline curvature technique which is used to compute the hub-to-tip flow field of a given turbomachine. To analyze the flow field in the blade-to-blade plane of the machine, the potential flow solution of an infinite cascade of airfoils is also computed using a source model technique. To verify the accuracy of such an analysis method an extensive experimental verification investigation was conducted using an axial flow research fan. Detailed surveys of the blade-free regions of the machine along with intra-blade surveys using rotating pressure sensing probes and blade surface static pressure taps provide a one-to-one relationship between measured and predicted data. The results of this investigation indicate the ability of this inviscid analysis method to predict the design flow field of the axial flow fan test rotor to within a few percent of the measured values.

Pierzga, M. J.↗
NASA NTRS2000

Heat Transfer and Pressure Distributions on a Gas Turbine Blade Tip

Heat transfer coefficient and static pressure distributions are experimentally investigated on a gas turbine blade tip in a five-bladed stationary linear cascade. The blade is a 2-dimensional model of a first stage gas turbine rotor blade with a blade tip profile of a GE-E(sup 3) aircraft gas turbine engine rotor blade. The flow condition in the test cascade corresponds to an overall pressure ratio of 1.32 and exit Reynolds number based on axial chord of 1.1 x 10(exp 6). The middle 3-blade has a variable tip gap clearance. All measurements are made at three different tip gap clearances of about 1%, 1.5%, and 2.5% of the blade span. Heat transfer measurements are also made at two different turbulence intensity levels of 6.1 % and 9.7% at the cascade inlet. Static pressure measurements are made in the mid-span and the near-tip regions as well as on the shroud surface, opposite the blade tip surface. Detailed heat transfer coefficient distributions on the plane tip surface are measured using a transient liquid crystal technique. Results show various regions of high and low heat transfer coefficient on the tip surface. Tip clearance has a significant influence on local tip beat transfer coefficient distribution. Heat transfer coefficient also increases about 15-20% along the leakage flow path at higher turbulence intensity level of 9.7% over 6.1 %.

Azad, Gm S.↗
NASA NTRS1973

An adverse effect of film cooling on the suction surface of a turbine vane

A J-75 size turbine vane with film cooling holes on the suction surface near the leading edge was tested with and without film cooling flow in a four vane cascade. Results show that the cooling effectiveness on the aft portion of the vane suction surface can decrease with the addition of film cooling near the leading edge. Apparently the film cooling air flow caused a laminar or transitional boundary layer to become a transitional or turbulent boundary layer. The vane was tested at a gas temperature and pressure of 1260 K (1800 F) and 22.7 newtons per square centimeter (33 psia), a coolant temperature of 280 K (50 F), film cooling flow ratios from 0.0 to 0.026, and backside midchord cooling flow ratios of 0.007 and 0.035.

Gladden, H. J.↗
NASA NTRS1985

A supersonic fan equipped variable cycle engine for a Mach 2.7 supersonic transport

The concept of a variable cycle turbofan engine with an axially supersonic fan stage as powerplant for a Mach 2.7 supersonic transport was evaluated. Quantitative cycle analysis was used to assess the effects of the fan inlet and blading efficiencies on engine performance. Thrust levels predicted by cycle analysis are shown to match the thrust requirements of a representative aircraft. Fan inlet geometry is discussed and it is shown that a fixed geometry conical spike will provide sufficient airflow throughout the operating regime. The supersonic fan considered consists of a single stage comprising a rotor and stator. The concept is similar in principle to a supersonic compressor, but differs by having a stator which removes swirl from the flow without producing a net rise in static pressure. Operating conditions peculiar to the axially supersonic fan are discussed. Geometry of rotor and stator cascades are presented which utilize a supersonic vortex flow distribution. Results of a 2-D CFD flow analysis of these cascades are presented. A simple estimate of passage losses was made using empirical methods.

Tavares, T. S.↗
NASA NTRS1980

Experimental determination of unsteady blade element aerodynamics in cascades. Volume 2: Translation mode cascade

A two dimensional cascade of harmonically oscillating airfoils was designed to model a near tip section from a rotor which was known to have experienced supersonic translational model flutter. This five bladed cascade had a solidity of 1.52 and a setting angle of 0.90 rad. Unique graphite epoxy airfoils were fabricated to achieve the realistic high reduced frequency level of 0.15. The cascade was tested over a range of static pressure ratios approximating the blade element operating conditions of the rotor along a constant speed line which penetrated the flutter boundary. The time steady and time unsteady flow field surrounding the center cascade airfoil were investigated.

Riffel, R. E.↗