A Systematic Computational Design System for Turbine Cascades, Airfoil Geometry and Blade-to-Blade Analysis

This paper describes a systematic computational design system for two-dimensional turbine cascades. The system includes a sequence of calculations in which airfoil profiles are designed from velocity diagram requirements and specified geometric parameters, followed by an inviscid global streamline curvature analysis, a magnified reanalysis around the leading edge, and a transitional profile boundary layer and wake mixing analysis. A finite area technique and a body-fitted mesh are used for the reanalysis. The boundary layer analysis is performed using the dissipation-integral method of Walz which has been modified in the present application. Several turbine airfoil profile geometry designs are presented. Also two sample cascade design cases and their calculated performance for a range of Mach numbers and incidence angles are given and discussed.

The Application of Adiabatic Film Cooling Effectiveness to Non-Adiabatic Blade Cooling Design

The effect of heat transfer on film cooling has been studied analytically. The proposed model shows that the non-adiabatic film cooling effectiveness will increase with increasing of the heat transfer parameter, Ū / (ρVCp)2, on the convex, the flat and the concave walls over the entire range of film cooling parameter, X/MS. On the convex wall with a blowing rate, M, of 0.51 and a heat transfer parameter of 10−3 at the typical engine conditions, the non-adiabatic effectiveness can be higher than the adiabatic effectiveness by 45% at a film cooling parameter of 103; while the film temperature can be lower than the adiabatic wall by 18°C (32°F) at a dimensionless distance of 500. The model can be extended and applied to the heat transfer analysis for any kind of turbine blade with film cooling.

Temperature and Dilution Effects in Soot Mass Concentration Measurements

Measurements are made of soot mass concentration in a luminous, liquid fuel spray, diffusion flame at atmospheric pressure. Intrusive sampling probes are used to study the effects of sampling rate, cooling, nitrogen-dilution ratio, and tip geometry on the mass of soot particles deposited on filters. Probe diameters have been kept small to minimize disturbance to the flow-field. Relative soot concentrations are observed to be lowest for uncooled probes, higher for water-cooled probes and still higher for probes with both water cooling and nitrogen injection. Furthermore, soot concentration steadily rises as the nitrogen/sample dilution ratio is increased from zero to as high as 1.5. Sampling rate has little effect on soot concentrations under most, but not all, sampling conditions.

Combined Cycle and Waste Heat Recovery Power Systems Based on a Novel Thermodynamic Energy Cycle Utilizing Low-Temperature Heat for Power Generation

A new thermodynamic energy cycle has been developed, using a multicomponent working agent. Condensation is supplemented with absorption, following expansion in the turbine. Several combined power systems based on this cycle have been designed and cost-estimated. Efficiencies of these new systems are 1.35 to 1.5 times higher than the best Rankine Cycle system, at the same border conditions. Investment cost per unit of power output is about two-thirds of the cost of a comparable Rankine Cycle system. Results make cogeneration economically attractive at current energy prices. The first experimental installation is planned by Fayette Manufacturing Company and Detroit Diesel Allison Division of General Motors.

Parametric Study of Flame Radiation Characteristics of a Tubular-Can Combustor

A series of combustor tests were conducted with a tubular-can combustor to study flame radiation characteristics and effects with parametric variations in combustor operating conditions. Two alternate combustor assemblies using a different fuel nozzle were compared. Spectral and total radiation detectors were positioned at three stations along the length of the combustor can. Data were obtained for a range of pressures from 0.34 to 2.07 MPa (50 to 300 psia), inlet temperatures from 533 to 700K (500 to 800°F), for Jet A (13.9% hydrogen) and ERBS (12.9% hydrogen) fuels, and with fuel-air ratios nominally from 0.008 to 0.021. Spectral radiation data, total radiant heat flux data, and liner temperature data are presented to illustrate the flame radiation characteristics and effects in the primary, secondary, and tertiary combustion zones.

A Comparison of Techniques for Measuring Particulate Emissions of a Gas Turbine Engine

The particulate emissions of an AVCO Lycoming vehicular gas turbine were measured by several techniques — two for determining smoke number and three for measurement of true smoke density. There was little agreement among the different techniques, with smoke numbers varying between 16 and 31 and true smoke density ranging from 6 to 19 milligrams per cubic meter at one engine test condition. The differences were attributed to timewise variation in engine performance, effects of different sampling techniques and the presence of volatile matter in the particulates. Particulate measurements were made during steady state and transient operation using a new instrument called a Tapered Element Oscillating Microbalance.

NASA Advanced Low Emissions Combustor Program

The purpose of this program is to conduct combustion tests on lean, premixed, and prevaporized (LPP) combustor concepts designed for use in commercial aircraft engines to attain improved performance, durability, and lower pollutant emissions levels relative to current technology combustor designs. Four full annular combustors were designed for the CF6–50 engine. These concepts utilize premixing of the fuel and air, variable geometry, and fuel staging to control the equivalence ratios of the burning zone. The testing is being conducted on these four full annular combustors over a wide range of operating conditions at pressures up to actual subsonic cruise (1.16 MPa). The test results for the most promising of these combustor concepts are reported in this paper.

Axial Fan Performance With Blade-Base Clearance

A study to evaluate the influence of increasing the clearance between blade and hub on a controllable pitch axial fan (CPAF) is presented. Fan performance was measured over a range of increasing clearance for several settings of blade pitch angles. The resulting variations of pressure rise, flow rate and efficiency have been correlated as functions of established clearance parameters with good results. The study shows that large base clearances may result in reductions in efficiency and flow rate of 5 percent or greater in a typical CPAF configuration.

The NASA Broad-Specification Fuels Combustion Technology Program: An Assessment of Phase I Test Results

An assessment is made of the results of phase I screening testing of current and advanced combustion system concepts using several broadened-properties fuels. The severity of each of several fuels-properties effects on combustor performance or liner life is discussed, as well as design techniques with the potential to offset these adverse effects. The selection of concepts to be pursued in Phase II refinement testing is described. This selection takes into account the relative costs and complexities of the concepts, the current outlook on pollutant emissions control, and practical operational problems.

A Variable-Geometry Combustor Used to Study Primary and Secondary Zone Stoichiometry

A combustion program is underway to evaluate fuel quality effects on gas turbine combustors. A rich-lean variable geometry combustor design was chosen to evaluate fuel quality effects over a wide range of primary and secondary zone equivalence ratios at simulated engine operating conditions. The first task of this effort, was to evaluate the performance of the variable geometry combustor. The combustor incorporates three stations of variable geometry to control primary and secondary zone equivalence ratio and overall pressure loss. Geometry changes could be made while a test was in progress through the use of remote control actuators. The primary zone liner was water cooled to eliminate the concern of liner durability. Emissions and performance data were obtained at simulated engine conditions of 80 percent and full power. Inlet air temperature varied from 611 to 665 K, inlet total pressure varied from 1.02 to 1.24 MPa, reference velocity was 18.0 m/sec and exhaust gas temperature was a constant 1400 K.