Fluid Flow and Heat Transfer Characteristics in the Entrance Regions of Circular Pipes

This paper describes the results of an experimental investigation into the velocity and turbulence fields and to a lesser extent the heat transfer in the entrance regions of short, circular cross-section pipes with length to diameter ratios up to 20 over the Reynolds number range from 35,000 to 170,000. The velocity and turbulence fields were measured by hot-wire anemometers backed up with pressure measurements and flow visualisation and the heat transfer by heat flux meters.

A Study of a Turbine-Generator System for Low-Head Hydropower

A method is outlined for determining the optimum operating conditions of a turbine-generator unit installed across a low-head irrigation structure for electrical power generation. For a given regulator’s characteristic, the unit’s rated power and design parameters are determined such that its cost-benefit ratio is minimum. The economical feasibility of the microhydro plant is studied by comparing its life-time cost to its life-time benefit. The benefit is determined by the cost of the corresponding energy generated through a diesel-driven generator set. The microhydro plant was found to be economically feasible over a wide range of inflation and interest rates.

Development of a Design Model for Airfoil Leading Edge Film Cooling

Leading edge showerhead cooling designs represent an important feature of certain classes of high temperature turbine airfoils. This paper outlines a methodology for predicting the surface temperatures of showerhead designs with spanwise injection through an array of discrete holes. The paper describes a series of experiments and analyses on scaled cylinder models with injection through holes inclined at 20, 30, 45, and 90 degrees for typical radial and circumferential spacing-to-diameter ratios of 10 and 4, respectively. The experiments were conducted in a wind tunnel on several stainless steel test specimens in which flow and heat transfer parameters were measured over the simulated airfoil leading edge surfaces. Based on the experiments, an engineering design model is proposed that treats the gas-to-surface heat transfer coefficient with film cooling in a manner suggested by a recent Purdue-NASA investigation and includes the important contribution of upstream (coolant inlet face) heat transfer. The experiments suggest that the averaged film cooling effectiveness in the showerhead region is primarily influenced by the inclination of the injection holes. The effectiveness parameter is not strongly affected by variations in coolant-to-gas stream pressure ratio, freestream Mach number, gas-to-coolant temperature ratio and gas stream Reynolds number. This is appropriately reflected in the design model in which the increase in coolant side heat transfer coefficient (with blowing ratio) is essentially offset by a simultaneous increase in the gas side film coefficient. The model is also employed to determine (inferentially) the average Stanton number reduction parameter for a series of pressure ratios varying from 1.004 to 1.3, Mach numbers ranging from 0.1 to 0.2, temperature ratios between 1.6 and 2.0, and Reynolds numbers ranging from 3.5 × 104 to 9.0 × 104. Design capabilities of the analytical model are explored for typical high temperature first stage turbine vanes and rotor blades operating at rotor inlet temperatures in excess of 1644°K.

Discharge Coefficients of Impingement and Film Cooling Holes

In this article measurements of fluid flow through impingement and film cooling holes for typical turbine blade cooling systems are presented. The purpose of the measurements was to determine hole discharge coefficients over a range of Reynolds numbers from 5,000 to 30,000 and to observe in this range the dependence of discharge coefficient on Reynolds number. The effect of hole geometry, that is, sharp edged inlet or corner radius inlet, on discharge coefficients is also measured. Correlations relating discharge coefficients to Reynolds number, corner radius to hole diameter ratio, and blowing parameter are suggested.

Full Coverage Discrete Hole Film Cooling: The Influence of Hole Size

The influence of hole size and hence blowing rate on full coverage discrete hole wall cooling for gas turbine combustion chamber applications was investigated. Two temperature conditions were used firstly a 750K gas temperature and 300K coolant, and secondly a realistic combustor primary zone conditions of 2100K flame temperature and 700K coolant. It was shown that a large hole size resulted in a significant improvement in the overall cooling effectiveness due to a reduced film heat transfer coefficient. At high temperature the cooling effectiveness was reduced due to radiative heat transfer from the flame gases. At low coolant flow large temperature increases of the coolant occurred within the wall and approached the transpiration situation.

Heat Transfer Measurements With Film Cooling on a Turbine Blade Profile in Cascade

Heat transfer measurements with film cooling have been made on a gas turbine rotor profile in a cascade at engine representative operating conditions. The blade temperature was varied independently to investigate the scaling of heat transfer coefficient, and a superposition model was found to correlate the data. Contrasting results are presented for films on the two surfaces, along with predictions from a 2–D boundary layer method.

Inverse Design of Coolant Flow Passage Shapes With Partially Fixed Internal Geometries

A method is described for the inverse design of complex coolant flow passage shapes in internally cooled turbine blades. This method is a refinement and extension of a method developed by the authors for designing a single coolant hole in turbine blades. The new method allows the turbine designer to specify the number of holes the turbine blade is to have. In addition, the turbine designer may specify that certain portions of the interior coolant flow passage geometry are to remain fixed (eg. struts, surface coolant ejection channels, etc.). Like the original design method, the designer must specify the outer blade surface temperature and heat flux distribution and the desired interior coolant flow passage surface temperature distributions. This solution procedure involves satisfying the dual Dirichlet and Neumann specified boundary conditions of temperature and heat flux on the outer boundary of the airfoil while iteratively modifying the shapes of the coolant flow passages using a least squares optimization procedure that minimizes the error in satisfying the specified Dirichlet temperature boundary condition on the surface of each of the evolving interior holes. Portions of the inner geometry that are specified to be fixed are not modified. A first order panel method is used to solve Laplace’s equation for the steady heat conduction within the solid portions of the hollow blade, making the inverse design procedure very efficient and applicable to realistic geometries. Results are presented for a realistic turbine blade design problem.

Development of the Centaur Type H Gas Turbine Engine

In response to the needs of the market place for turbines in the 5000 to 6000 hp class, Solar Turbines Incorporated has responded with an uprate of their Centaur engine. Discussed in this paper are the features of the uprated engine, the Development Plan and the methodology for incorporating into the design the advanced aerodynamic and mechanical technology of the Mars engine. The Mars engine is a high efficiency 12,500 hp engine which operates at a turbine inlet temperature of 1935°F. State-of-the-art computer aided methods have been applied to produce the design, and the results from this approach are displayed.

Heat Transfer Results and Operational Characteristics of the NASA Lewis Research Center Hot Section Cascade Test Facility

The NASA Lewis Research Center gas turbine hot section test facility has been developed to provide a “real-engine” environment with well known boundary conditions for the aerothermal performance evaluation/verification of computer design codes. The initial aerothermal research data obtained at this facility are presented and the operational characteristics of the facility are discussed. This facility is capable of testing at temperatures and pressures up to 1600 K and 18 atm which corresponds to a vane exit Reynolds number range of 0.5×106 to 2.5×106 based on vane chord. The component cooling air temperature can be independently modulated between 330 and 700 K providing gas-to-coolant temperature ratios similar to current engine application. Research instrumentation of the test components provide conventional pressure and temperature measurements as well as metal temperatures measured by IR-photography. The primary data acquisition mode is steady state through a 704 channel multiplexer/digitizer. The test facility was configured as an annular cascade of full coverage film cooled vanes for the initial series of research tests. These vanes were tested over a wide range of gas Reynolds number, exit gas Mach number and heat flux levels. The range of test conditions was used to represent both actual operating conditions and similarity state conditions of a gas turbine engine. The results are presented for the aerothermal performance of the facility and the full coverage film cooled vanes.

A Review and Analysis of Boundary Layer Transition Data for Turbine Application

A Symposium on Transition in Turbines was held recently 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. This paper represents a preliminary response to that recommendation. 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 were analyzed 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 tables 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.