Engineering topics

Gaugler, R. E.

Publications and source records attributed to Gaugler, R. E..

At least 19 records

NASA NTRS2001

Two-Equation Turbulence Models for Prediction of Heat Transfer on a Transonic Turbine Blade

Two versions of the two-equation k-omega model and a shear stress transport (SST) model are used in a three-dimensional, multi-block, Navier-Stokes code to compare the detailed heat transfer measurements on a transonic turbine blade. It is found that the SST model resolves the passage vortex better on the suction side of the blade, thus yielding a better comparison with the experimental data than either of the k-w models. However, the comparison is still deficient on the suction side of the blade. Use of the SST model does require the computation of distance from a wall, which for a multiblock grid, such as in the present case, can be complicated. However, a relatively easy fix for this problem was devised. Also addressed are issues such as (1) computation of the production term in the turbulence equations for aerodynamic applications, and (2) the relation between the computational and experimental values for the turbulence length scale, and its influence on the passage vortex on the suction side of the turbine blade.

Garg, Vijay K.↗

NASA NTRS2000

Modeling Film-Coolant Flow Characteristics at the Exit of Shower-Head Holes

The coolant flow characteristics at the hole exits of a film-cooled blade are derived from an earlier analysis where the hole pipes and coolant plenum were also discretized. The blade chosen is the VKI rotor with three staggered rows of shower-head holes. The present analysis applies these flow characteristics at the shower-head hole exits. A multi-block three-dimensional Navier-Stokes code with Wilcox's k-omega model is used to compute the heat transfer coefficient on the film-cooled turbine blade. A reasonably good comparison with the experimental data as well as with the more complete earlier analysis where the hole pipes and coolant plenum were also gridded is obtained. If the 1/7th power law is assumed for the coolant flow characteristics at the hole exits, considerable differences in the heat transfer coefficient on the blade surface, specially in the leading-edge region, are observed even though the span-averaged values of h (heat transfer coefficient based on T(sub o)-T(sub w)) match well with the experimental data. This calls for span-resolved experimental data near film-cooling holes on a blade for better validation of the code.

Garg, Vijay K.↗

NASA NTRS1997

Effect of Velocity and Temperature Distribution at the Hole Exit on Film Cooling of Turbine Blades

An existing three-dimensional Navier-Stokes code (Arnone et al, 1991), modified Turbine Branch, to include film cooling considerations (Garg and Gaugler, 1994), has been used to study the effect of coolant velocity and temperature distribution at the hole exit on the heat transfer coefficient on three film-cooled turbine blades, namely, the C3X vane, the VKI rotor, and the ACE rotor. Results are also compared with the experimental data for all the blades. Moreover, Mayle's transition criterion (1991), Forest's model for augmentation of leading edge heat transfer due to free-stream turbulence (1977), and Crawford's model for augmentation of eddy viscosity due to film cooling (Crawford et al, 1980) are used. Use of Mayle's and Forest's models is relevant only for the ACE rotor due to the absence of showerhead cooling on this rotor. It is found that, in some cases, the effect of distribution of coolant velocity and temperature at the hole exit can be as much as 60 percent on the heat transfer coefficient at the blade suction surface, and 50 percent at the pressure surface. Also, different effects are observed on the pressure and suction surface depending upon the blade as well as upon the hole shape, conical or cylindrical.

Garg, V. K.↗

NASA NTRS1985

A review and analysis of boundary layer transition data for turbine application

A number of data sets from the open literature that include heat transfer data in apparently transitional boundary layers, with particular application to the turbine environment, were reviewed and analyzed to extract transition information. The data were analyzed by using a version of the STAN5 two-dimensional boundary layer code. The transition starting and ending points were determined by adjusting parameters in STAN5 until the calculations matched the data. The results are presented as a table of the deduced transition location and length as functions of the test parameters. The data sets reviewed cover a wide range of flow conditions, from low-speed, flat-plate tests to full-scale turbine airfoils operating at simulated turbine engine conditions. The results indicate that free-stream turbulence and pressure gradient have strong, and opposite, effects on the location of the start of transition and on the length of the transition zone.

Gaugler, R. E.↗

NASA NTRS1985

A review and analysis of boundary layer transition data for turbine application

A symposium on transition in turbines was held at the NASA Lewis Research Center. One recommendation of the working groups was the collection of existing transition data to provide standard cases against which models could be tested. A number of data sets from the open literature that include heat transfer data in apparently transitional boundary layers, with particular application to the turbine environment, were reviewed and analyzed to extract transition information from the heat transfer data. The data sets reviewed cover a wide range of flow conditions, from low speed, flat plate tests to full scale turbine airfoils operating at simulated turbine engine conditions. The results indicate that free stream turbulence and pressure gradient have strong, and opposite, effects on the location of the start of transition and on the length of the transition zone.

Gaugler, R. E.↗

NASA NTRS1984

Comparison of visualized turbine endwall secondary flows and measured heat transfer patterns

Various flow visualization techniques were used to define the seondary flows near the endwall in a large heat transfer data. A comparison of the visualized flow patterns and the measured Stanton number distribution was made for cases where the inlet Reynolds number and exit Mach number were matched. Flows were visualized by using neutrally buoyant helium-filled soap bubbles, by using smoke from oil soaked cigars, and by a few techniques using permanent marker pen ink dots and synthetic wintergreen oil. Details of the horseshoe vortex and secondary flows can be directly compared with heat transfer distribution. Near the cascade entrance there is an obvious correlation between the two sets of data, but well into the passage the effect of secondary flow is not as obvious. Previously announced in STAR as N83-14435

Gaugler, R. E.↗

NASA NTRS1984

A review and analysis of boundary layer transition data for turbine application

Gaugler, R. E.↗

NASA NTRS1983

Comparison of predicted and experimental external heat transfer around a film cooled cylinder in crossflow

Calculations were made of the film cooling provided by rows of holes around the circumference of a cylinder in crossflow and the results were compared to experimental data. The calculations and experimental data were for conditions that simulate most of those that are typical of air cooled turbine vane leading edges. Injection was from a single and multiple rows of holes located at different angular locations from the stagnation line. The holes in the rows were angled normal to the flow direction and at a 25 degree angle to the cylinder wall. The calculations and experimental data were for several constant values of blowing ratios for all rows and for different blowing ratios for each row, representing a simulation of a common coolant plenum supply to multiple rows of holes. The calculations were made using a finite difference boundary layer code, STAN5. Contrary to initial expectations that injection would trip the boundary layer flow into the turbulent regime, the results indicated that the high free stream acceleration apparently kept the flow laminar for holes in the first 45 degrees past stagnation. The trend in Stanton number reduction due to coolant injection was predicted with generally good agreement at the lower blowing rates, but for multile rows of holes, agreement was poor beyond the first row.

Stepka, F. S.↗

NASA NTRS1983

Comparison of visualized turbine endwall secondary flows and measured heat transfer patterns

Various flow visualization techniques were used to define the secondary flows near the endwall in a large heat transfer data. A comparison of the visualized flow patterns and the measured Stanton number distribution was made for cases where the inlet Reynolds number and exit Mach number were matched. Flows were visualized by using neutrally buoyant helium-filled soap bubbles, by using smoke from oil soaked cigars, and by a few techniques using permanent marker pen ink dots and synthetic wintergreen oil. Details of the horseshoe vortex and secondary flows can be directly compared with heat transfer distribution. Near the cascade entrance there is an obvious correlation between the two sets of data, but well into the passage the effect of secondary flow is not as obvious.

Gaugler, R. E.↗

NASA NTRS1982

Flow visualization study of the horseshoe vortex in a turbine stator cascade

Flow visualization techniques were used to show the behavior of the horseshoe vortex in a large scale turbine stator cascade. Oil drops on the end wall surface flowed in response to local shear stresses, indicating the limiting flow streamlines at the surface. Smoke injected into the flow and photographed showed time averaged flow behavior. Neutrally bouyant helium filled soap bubbles followed the flow and showed up on photographs as streaks, indicating the paths followed by individual fluid particles. Preliminary attempts to control the vortex were made by injecting air through control jets drilled in the end wall near the vane leading edge. Seventeen different hole locations were tested, one at a time, and the effect of the control jets on the path follwed by smoke in the boundary layer was recorded photographically.

Gaugler, R. E.↗

NASA NTRS1982

Measured and calculated wall temperatures on air-cooled turbine vanes with boundary layer transition

Convection cooled turbine vane metal wall temperatures experimentally obtained in a hot cascade for one vane design were compared with wall temperatures calculated with TACT1 and STAN5 computer codes which incorporated various models for predicting laminar-to-turbulent boundary layer transition. Favorable comparisons on both vane surface were obtained at high Reynolds number with only one of these transition models. When other models were used, temperature differences between calculated and experimental data obtained at the high Reynolds number were as much as 14 percent in the separation bubble region of the pressure surface. On the suction surface and at lower Reynolds number, predictions and data unsatisfactorily differed by as much as 22 percent. Temperature differences of this magnitude can represent orders of magnitude error in blade life prediction.

Liebert, C. H.↗

NASA NTRS1981

Some modifications to, and operational experiences with, the two-dimensional, finite-difference, boundary-layer code, STAN5

The two-dimensional, finite-difference boundary-layer code, STAN5, is the primary tool used at the NASA-Lewis Research Center for predicting turbine blade gas-side heat-transfer coefficients. A number of modifications have been made to the program to enhance its usefulness for these calculations. Experience in using STAN5 has identified some problems in the program that can be treated through program input, without modifying the program. These include the presence of a small separation bubble near the leading edge, and the effect of full-coverage film cooling on transition to turbulence. Some of the techniques used to treat these problems are described.

Gaugler, R. E.↗

NASA NTRS1980

Nonlinear, three-dimensional finite-element analysis of air-cooled gas turbine blades

Cyclic stress-strain states in cooled turbine blades were calculated for a simulated mission of an advanced-technology commercial aircraft engine. The MARC, nonlinear, finite-element computer program was used for the analysis of impingement-cooled airfoils, with and without leading-edge film cooling. Creep was the predominant damage mode (ignoring hot corrosion), particularly artund film-cooling holes. Radially angled holes exhibited less creep than holes with axes normal to the surface. Beam-theory analyses of all-impingement-cooled airfoils gave fair agreement with MARC results for initial creep.

Kaufman, A.↗

NASA NTRS1980

Streakline flow visualization study of a horseshoe vortex in a large-scale, two-dimensional turbine stator cascade

Neutrally buoyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large-scale, two-dimensional, turbine stator cascade. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16-mm movie film. Individual frames from the film have been selected, and overlayed to show the details of the horseshoe vortex around the leading edge. The transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

Gaugler, R. E.↗

NASA NTRS1980

Significance of thermal contact resistance in two-layer thermal-barrier-coated turbine vanes

The importance of thermal contact resistance between layers in heat transfer through two layer, plasma sprayed, thermal barrier coatings applied to turbine vanes was investigated. Results obtained with a system of NiCrAlY bond and yttria stabilized zirconia ceramic show that thermal contact resistance between layers is negligible. These results also verified other studies which showed that thermal contact resistance is negligible for a different coating system of NiCr bond calcia stabilized zirconia ceramic. The zirconia stabilized ceramic thermal conductivity data scatter presented in the literature is ?20 to -10 percent about a curve fit of the data. More accurate predictions of heat transfer and metal wall temperatures are obtained when the thermal conductivity values are used at the ?20 percent level.

Liebert, C. H.↗

NASA NTRS1979

Thermal-structural mission analyses of air-cooled gas turbine blades

Cyclic temperature and stress-strain states in cooled turbine blades were calculated for a simulated mission of an advanced technology aircraft engine. TACTl (three dimensional heat transfer) and MARC (nonlinear structural analysis) computer programs were used to analyze impingement cooled airfoils, with and without leading-edge film cooling. Creep was the predominant damage mode, particularly around film cooling holes. Radially angled holes exhibited less creep than holes normal to surface. Beam-type analyses of all-impingement cooled airfoils gave fair agreement with MARC results for initial creep.

Kaufman, A.↗

NASA NTRS1979

TACT 1: A computer program for the transient thermal analysis of a cooled turbine blade or vane equipped with a coolant insert. 2. Programmers manual

A computer program to calculate transient and steady state temperatures, pressures, and coolant flows in a cooled axial flow turbine blade or vane with an impingement insert is described. Coolant-side heat transfer coefficients are calculated internally in the program, with the user specifying either impingement or convection heat transfer at each internal flow station. Spent impingement air flows in a chordwise direction and is discharged through the trailing edge and through film cooling holes. The ability of the program to handle film cooling is limited by the internal flow model. Input to the program includes a description of the blade geometry, coolant-supply conditions, outside thermal boundary conditions, and wheel speed. The blade wall can have two layers of different materials, such as a ceramic thermal barrier coating over a metallic substrate. Program output includes the temperature at each node, the coolant pressures and flow rates, and the coolant-side heat transfer coefficients.

Gaugler, R. E.↗

NASA NTRS1979

Streakline flow visualization study of a horseshoe vortex in a large-scale, two-dimensional turbine stator cascade

Neutrally bouyant helium-filled bubbles were observed as they followed the streamlines in a horseshoe vortex system around the vane leading edge in a large scale, two dimensional, turbine stator cascade. Inlet Reynolds number, based on true chord, ranged between 100,000 to 300,000. Bubbles were introduced into the endwall boundary layer through a slot upstream of the vane leading edge. The paths of the bubbles were recorded photographically as streaklines on 16 mm movie film. Individual frames from the film were selected, and overlayed to show the details of the horseshoe vortex around the leading edge. The transport of the vortex across the passage near the leading edge is clearly seen when compared to the streaks formed by bubbles carried in the main stream. Limiting streamlines on the endwall surface were traced by the flow of oil drops.

Gaugler, R. E.↗