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NASA NTRS1993

Three-dimensional flow field in a turbine nozzle passage

Detailed measurements were taken in the nozzle of a low speed single stage axial flow turbine at two axial planes inside the nozzle and on the nozzle and endwall surfaces. Velocity, turbulence and angle measurements were taken at midchord with an LDV while a five hole probe was used to measure the pressure, velocities and angles just upstream of the trailing edge. Nozzle surface and endwall static pressures were also measured. These measurements were compared to measurements previously completed at two axial planes downstream of the nozzle. The results show that at midchord, the secondary flow seems to be weak and it is in the early stages of development. Just upstream of the trailing edge, the secondary flow is clearly visible. The radially inward flow near the suction surface augments the casing passage vortex, while counteracting the hub passage vortex. Traveling downstream, the casing passage vortex remains strong while at the hub, the radially inward flow of the suction surface boundary layer has reversed direction due to the rotating hub. The blade static pressures and the passage averaged velocities compare well with Katsanis' quasi-three-dimensional code. These and other data are presented, interpreted and synthesized to understand the nozzle flow field.

Zaccaria, M.↗

NASA NTRS1980

A comparison between an existing propeller noise theory and wind tunnel data

The noise of three supersonic helical tip speed propellers was compared with the noise predicted by an existing noise theory. Comparisons of the peak blade passage tones showed fairly good agreement beween theory and experiment at the lowest helical tip Mach numbers tested, 0.86 and 1.00, while at higher numbers, the theory predicted higher noise levels than measured. When the differences among the propellers were considered the theory and measurement showed fairly good agreement. Directivity measurements in general showed that the measured blade passage tone data peaked further downstream than the theory predicted. At the cruise design condition the harmonics appeared to fall off faster in the data than the theory indicated.

Dittmar, J. H.↗

NASA NTRS1989

Plasma density, temperature and turbulence in the wake of the Shuttle Orbiter

In situ observations of the plasma density, temperature, and turbulence in the near wake, midwake, and far wake of the Shuttle Orbiter are presented. The results show that the plasma disturbances produced by the passage of the Orbiter extend downstream for a distance of several hundred meters behind the Orbiter. It is found that the electron density profile exhibits nonmonotonic behavior, and that most of the 6-40-Hz turbulence in the wake is confined to a region about 20 m downstream of the Orbiter. The saturation Delta N(e)/N(e) spectra rapidly decline in the 10,000-100,000-Hz decade, both in the ambient ionosphere and in the wake.

Tribble, A. C.↗

NASA NTRS1975

Effect of wall edge suction on the performance of a short annular dump diffuser with exit passage flow resistance

The effect of wall edge suction on the performance of a short annular dump diffuser having a perforated plate flow resistance device in the exit passage was evaluated. Testing was conducted with air at near ambient pressure and temperature at inlet Mach numbers of 0.18 and 0.27 with suction rates up to 13.5 percent. Results show that pressure recovery downstream of the perforated plate was improved significantly by suction. Optimum performance was obtained with the flow resistance plate located at one inlet passage height downstream of the dump plane.

Juhasz, A. J.↗

NASA NTRS1981

Characteristics of mean velocity in the annulus wall region at the exit of turbomachinery rotor passage

This paper reports the experimental study of the three-dimensional characteristics of the mean velocity in the annulus-wall boundary layer and wake mixing region of a moderately loaded compressor rotor blade. The measurements reported were obtained using a three-sensor hot-wire probe at the exit of the rotor. They consisted of tangential surveys at several axial and radial locations downstream of the compressor rotor blade passage. The variation of the streamwise, normal, and radial velocity profiles and their decay characteristics are derived from this data. The measurements reveal the highly complex and three-dimensional nature of the annulus-wall flow field and the nature of the interaction of the annulus-wall boundary layer, the blade boundary layers and wake, and the secondary flow associated with the annulus-wall region.

Davino, R.↗

NASA NTRS1993

Heat transfer in rotating serpentine passages with selected model orientation for smooth or skewed trip walls

Experiments were conducted to determine the effects of model orientation as well as buoyancy and Coriolis forces on heat transfer in turbine blade internal coolant passages. Turbine blades have internal coolant passage surfaces at the leading and trailing edges of the airfoil with surfaces at angles which are as large as +/- 50 to 60 degrees to the axis of rotation. Most of the previously-presented, multiple-passage, rotating heat transfer experiments have focused on radial passages aligned with the axis of rotation. Results from serpentine passages with orientations 0 and 45 degrees to the axis of rotation which simulate the coolant passages for the mid chord and trailing edge regions of the rotating airfoil are compared. The experiments were conducted with rotation in both directions to simulate serpentine coolant passages with the rearward flow of coolant or with the forward flow of coolant. The experiments were conducted for passages with smooth surfaces and with 45 degree trips adjacent to airfoil surfaces for the radial portion of the serpentine passages. At a typical flow condition, the heat transfer on the leading surfaces for flow outward in the first passage with smooth walls was twice as much for the model at 45 degrees compared to the model at 0 degrees. However, the differences for the other passages and with trips were less. In addition, the effects of buoyancy and Coriolis forces on heat transfer in the rotating passage were decreased with the model at 45 degrees, compared to the results at 0 degrees. The heat transfer in the turn regions and immediately downstream of the turns in the second passage with flow inward and in the third passage with flow outward was also a function of model orientation with differences as large as 40 to 50 percent occurring between the model orientations with forward flow and rearward flow of coolant.

Johnson, B. V.↗

NASA NTRS1987

Laser Doppler velocimeter measurement of annulus wall boundary layer development in a compressor rotor

Detailed measurement of the flow field in the tip region of a compressor rotor was carried out using LDV. The axial and tangential components of relative velocities were measured upstream, inside the passage, and at the exit of the rotor, up to about 20 percent of the blade span from the blade tip. The annulus-wall boundary layer is well behaved at the leading edge and far downstream of the rotor. But inside the passage, complex interactions between the leakage flow and the annulus-wall boundary layer result in unconventional profiles with wide deviations from models employed for analyses.

Lakshminarayana, B.↗

NASA NTRS1982

Laser anemometer measurements in an annular cascade of core turbine vanes and comparison with theory

Laser measurements were made in an annular cascade of stator vanes operating at an exit critical velocity ratio of 0.78. Velocity and flow angles in the blade to blade plane were obtained at every 10 percent of axial chord within the passage and at 1/2 axial chord downstream of the vanes for radial positions near the hub, mean and tip. Results are presented in both plot and tabulated form and are compared with calculations from an inviscid, quasi three dimensional computer program. The experimental measurements generally agreed well with these theoretical calculations, an indication of the usefulness of this analytic approach.

Goldman, L. J.↗

NASA NTRS2001

Coupled-Flow Simulation of HP-LP Turbines Has Resulted in Significant Fuel Savings

Our objective was to create a high-fidelity Navier-Stokes computer simulation of the flow through the turbines of a modern high-bypass-ratio turbofan engine. The simulation would have to capture the aerodynamic interactions between closely coupled high- and low-pressure turbines. A computer simulation of the flow in the GE90 turbofan engine's high-pressure (HP) and low-pressure (LP) turbines was created at GE Aircraft Engines under contract with the NASA Glenn Research Center. The three-dimensional steady-state computer simulation was performed using Glenn's average-passage approach named APNASA. The areas upstream and downstream of each blade row mutually interact with each other during engine operation. The embedded blade row operating conditions are modeled since the average passage equations in APNASA actively include the effects of the adjacent blade rows. The turbine airfoils, platforms, and casing are actively cooled by compressor bleed air. Hot gas leaks around the tips of rotors through labyrinth seals. The flow exiting the high work HP turbines is partially transonic and, therefore, has a strong shock system in the transition region. The simulation was done using 121 processors of a Silicon Graphics Origin 2000 (NAS 02K) cluster at the NASA Ames Research Center, with a parallel efficiency of 87 percent in 15 hr. The typical average-passage analysis mesh size per blade row was 280 by 45 by 55, or approx.700,000 grid points. The total number of blade rows was 18 for a combined HP and LP turbine system including the struts in the transition duct and exit guide vane, which contain 12.6 million grid points. Design cycle turnaround time requirements ran typically from 24 to 48 hr of wall clock time. The number of iterations for convergence was 10,000 at 8.03x10(exp -5) sec/iteration/grid point (NAS O2K). Parallel processing by up to 40 processors is required to meet the design cycle time constraints. This is the first-ever flow simulation of an HP and LP turbine. In addition, it includes the struts in the transition duct and exit guide vanes.

Veres, Joseph P.↗

NASA NTRS1976

Three-dimensional flow calculation for a transonic compressor rotor

A numerical calculation of the steady, inviscid, three dimensional flow in a isolated transonic compressor rotor has been completed using MacCormack's second order accurate time-marching scheme. This rotor has a tip Mach number of 1.2, an overall diameter of 2 feet, and inlet hub/tip ratio of 0.5. The computed rotor total pressure ratio is 1.82. Comparisons between the numerical solution, measurements of the intra-blade static density field obtained by gas fluorescence, and time resolved exit flow measurements showed that the inviscid computation accurately models transonic rotor aerodynamics and rotor blade pressure distributions in the upstream portions of the blade passages, the viscous effects influencing mainly the downstream flow.

Thompkins, W. T., Jr.↗

NASA NTRS1991

Continuous wave laser absorption techniques for gasdynamic measurements in supersonic flows

Line-of-sight measurements of velocity, temperature, pressure, density, and mass flux were performed in a transient shock tube flow using three laser absorption schemes. All methods employed an intracavity-doubled ring dye laser tuned to an OH transition at 306 nm. In the first scheme, the gas was labeled by 193.3-nm excimer photolysis of H2O, and the passage of the generated OH was detected downstream. In the second method, the laser was tuned at a rate of 3 kHz over the R1(7) and R1(11) line pair, and absorption was simultaneously monitored at 90 and 60 deg with respect to the flow. Velocity was determined from the Doppler shift of the profiles and the temperature from the intensity ratio of the lines. Pressure was determined from both the magnitude of absorption and the collisional broadening. In the third method, the laser wavelength was fixed at a single frequency, and a continuous measurement of velocity and pressure was obtained using the signals from the two beam paths. All methods gave results which compare favorably to calculated values.

Davidson, David F.↗

NASA NTRS1990

Euler analysis comparison with LDV data for an advanced counter-rotation propfan at cruise

A fine mesh Euler solution of the F4/A4 unducted fan (UDF) model flowfield is compared with laser Doppler velocimeter (LDV) data taken in the NASA Lewis 8- by 6-Foot Supersonic Wind Tunnel. The comparison is made primarily at one axial plane downstream of the front rotor where the LDV particle lag errors are reduced. The agreement between measured and predicted velocities in this axial plane is good. The results show that a dense mesh is needed in the centerbody stagnation region to minimize entropy generation that weakens the aft row passage shock. The predicted radial location of the tip vortex downstream of the front rotor agrees well with the experimental results but the strength is overpredicted. With 40 points per chord line, the integrated performance quantities are nearly converged, but more points are needed to resolve passage shocks and flow field details.

Miller, Christopher J.↗

NASA NTRS1990

Euler analysis comparison with LDV data for an advanced counter-rotation propfan at cruise

Miller, Christopher J.↗

NASA NTRS1977

Flow in a transonic compressor rotor

Some results of a comprehensive study of the flow through a transonic compressor rotor are presented. The rotor produces a pressure ratio of 1.6 at a tip Mach number of 1.2, has a cylindrical casing, sloped hub, and MCA blade sections. The gas density has been measured in the rotor by gas fluorescence, which reveals details of the shock and boundary layer structure as well as density in the passages. Time resolved measurements of the flow just downstream of the rotor give the blade wake structure, which is related to the features shown by the fluorescence. Of particular interest is a complex three-dimensional separation near the sonic radius. The flow through the rotor has also been computed in three dimensions by a time-marching technique, and the results of this computation are compared to those found by flow visualization.

Kerrebrock, J. L.↗

NASA NTRS1972

Methods for reducing blade passing frequency noise generated by rotor-wake - stator interaction

A theoretical model for the generation of blade-passage noise caused by the interaction of rotor wakes with the downstream stator blades in a fan stage has been compiled. This model has combined a description of the rotor wake based on existing experimental correlations with an existing theory for the response of the stator blades to those wakes. This model of the interaction-caused blade-passage noise has been used to examine some effects of blade geometry and flow conditions on the noise generated. A set of significant parameters has been identified to reduce this interaction-caused blade-passage frequency noise. A redesign of an existing fan with the use of these concepts has produced a theoretical noise reduction of more than 7 decibels.

Dittmar, J. H.↗

NASA NTRS1974

Ordered structures and jet noise

A series of measurements of near field pressures and turbulent velocity fluctuations were made in a jet having a Reynolds number of about 50,000 in order to investigate more quantitatively the character and behavior of the large scale structures, and to ascertain their importance to the jet noise problem. It was found that the process of interaction between vortices can be inhibited by artificially exciting the shear layers with periodic disturbances of certain frequency. The turbulent fluctuation amplitudes measured at four diameters downstream decreased considerably. Finally, it was observed that the passage frequency of the structures decreased with x in a similar manner as the frequency corresponding to the maximum intensity radiation emanating from the same value of x.

Petersen, R. A.↗

NASA NTRS1992

Optimum design of ninety degree bends

An algorithm for the optimum design of an internal flow component to obtain the maximum pressure rise is presented. Maximum pressure rise in a duct with simultaneous turning and diffusion is shown to be related to the control of flow separation on the passage walls. Such a flow is usually associated with downstream conditions that are desirable in turbomachinery and propulsion applications to ensure low loss and stable performance. The algorithm requires the solution of an 'adjoint' problem in addition to the 'direct' equations governing the flow in a body, which in the present analysis are assumed to be the laminar Navier-Stokes equations. The theoretical framework and computational algorithms presented in this study are for the steady Navier-Stokes equations. A procedure is developed for the numerical solution of the adjoint equations. This procedure is coupled with a direct solver in a design iteration loop, that provides a new shape with a higher pressure rise. This procedure is first validated for the design of optimum plane diffusers in two-dimensional flow. The direct Navier-Stokes and the 'adjoint' equations are solved using a finite volume formulation for spatial discretization in an artificial compressibility framework. A simplified version of the above approach is then utilized to design ninety degree diffusing bends. Calculations were carried out for a mean radius ratio at inlet of 2.5 and Reynolds numbers varying from 100 to 500. While at this stage laminar flows is assumed, it is shown that a similar approach can be conceived for turbulent flows.

Modi, Vijay↗

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.↗