Improving Turbine Component Efficiency

Transcription

1 Copyright 1979 by ASME Improving Turbine Component Efficiency K. D. MACH AFAPL/TBC, Wright-Patterson Air Force Base, Ohio The thermal efficiency of a gas turbine engine depends on the cycle pressure and temperature ratio and on the aerodynamic efficiencies of the gas path components. Maintaining and/or improving structural integrity and aerodynamic efficiency in this high pressure, high temperature environment is the preeminent problem of the turbine designer. High gas temperatures require at least some of the metal structures to be cooled, yet cooling air is a loss to the cycle and its consumption must be kept to a minimum. Research into cooling techniques and boundary layer behavior on airfoils and endwalls and into test procedures tor obtaining heat transfer data are providing some of the answers the designer needs. Increased operating pressures generate increased mechanical stresses. Finite element analyses and automated design procedures are proving to be powerful aids to the designer. Improving aerodynamic efficiency requires careful control of the flow in three dimensions, particularly in low aspect ratio machines. The first practical computation method for three-dimensional, viscous, transonic flows became available in late 1977 and has made this one of the most exciting areas of turbine technology. Additional gains in aerodynamic efficiency can be realized by controlling leakages, especially those over the rotor tip, by accounting for the transient interactions between rotor and stator and by careful control of discharged coolant flow. This paper briefly describes the turbine cooling research conducted by the Air Force Aero Propulsion Laboratory and describes mor extensively the AFAPL programs in turbine aerodynamics, including applications of three-dimensional flow analysis. Contributed by the Gas Turbine Division of The American Society of Mechanical Engineers for presentation at the Gas Turbine Conference & Exhibit & Solar Energy Conference, San Diego, Calif., March 12-15, Manuscript received at ASME Headquarters January 8, Copies will be available until December 1, UNITED ENGINEERING CENTER, 345 EAST 47th STREET, NEW YORK, N.Y

2 Improving Turbine Component Efficiency K. D. MACH ABSTRACT The thermal efficiency of a gas turbine engine depends on the cycle pressure and temperature ratio and on the aerodynamic efficiencies of the gas path components. Maintaining and/or improving structural integrity and aerodynamic efficiency in this high pressure, high temperature environment is the preeminent problem of the turbine designer. High gas temperatures require at least some of the metal structure to be cooled, yet cooling air is a loss to the cycle and its consumption must be kept to a minimum. Research into cooling techniques and boundary layer behavior on airfoils and endwalls and into test procedures for obtaining heat transfer data are providing some of the answers the designer needs. Increased operating pressures generate increased mechanical stresses. Finite element analyses and automated design procedures are proving to be powerful aids to the designer. Improving aerodynamic efficiency requires careful control of the flow in three dimensions, particularly in low aspect ratio machines. The first practical computation method for three-dimensional, viscous, transonic flows became available in late 1977 and has made this one of the most exciting areas of turbine technology. Additional gains in aerodyanmic efficiency can be realized by controlling leakages, especially those over the rotor tip, by accounting for the transient interactions between rotor and stator; and by careful control of discharged coolant flow. This paper briefly describes the turbine cooling research conducted by the Air Force Aero Propulsion Laboratory and describes more extensively the AFAPL programs in turbine aerodynamics, including applications of three-dimensional flow analysis INTRODUCTION The thermal efficiency and specific output of a gas turbine engine depend on the cycle pressure and temperature ratios and on the aerodynamic efficiencies of the gas path components. Cycle efficiency increases with compressor pressure ratio and specific output increases with the ratio of turbine inlet temperature to compressor inlet temperature, hence the continuous striving for higher compressor pressure ratios and higher turbine inlet temperatures. Aerodynamic inefficiencies in the gas path degrade the cycle efficiency on at least a one-to-one ratio. For example, in a simple turbojet with compressor pressure ratio of 20:1 and temperature ratio (turbine inlet to compressor inlet) of 4.0, and all gas path efficiencies set to unity, the cycle efficiency is 57.5%. If we decrease the turbine efficiency, the cycle efficiency decreases 1.4 times as fast, reaching zero when the turbine efficiency is 59%. The efficiencies of other gas path components exert similar, if not greater, influences. It is clear, therefore, that the achievement of maximum component efficiency is imperative. In order to design an efficient turbine, today's designer must contend with three major interrelating problems. He must provide an efficient three-dimensional flow path, and he must provide some means of blade and endwall cooling, at least in the first stages, while accounting for the effect of discharged coolant on the main flow. Finally, he must provide 1

3 sufficient structural strength to withstand the high pressures and temperatures, allowing space for internal coolant flow, and holding weight to a minimum. Succeeding paragraphs of this paper will examine some of these problems in more detail and describe some of the research activities directed toward their solution. The paper concludes with a projection of the state-of-the-art of turbine design in the 1980's. DESCRIPTION Broadly speaking, the turbine design process consists of a cycle of preliminary aerodynamic and stress analyses followed by detailed and elaborate analyses. The preliminary analyses are iterative and serve to delimit the various design parameters-the flow path, the vector diagrams, the airfoil sections, the heat load on the airfoils, the required coolant, and the stress levels. Preliminary Design The actual nature of preliminary design and analysis calculations has not changed significantly over the years because the governing principles and the steps of their implementation have long been known. What has changed is the method of executing the solution procedure and to some extent the magnitude of the problem. The digital computer afforded a significant increase in designer productivity because of its speed. Problems which took days with a slide rule or desk calculator could be solved in minutes with-assuming a thoroughly debugged program-no errors. The introduction of interactive computer graphics yielded another major increase in designer productivity because he could obtain both graphic and printed output almost instantly, modify the input, and again see the result almost instantly. A preliminary design system of this type is described in the next section, and with greater detail in Reference 6. With it, a designer can analyze a complete stage-root, pitch, and tip sections-from vector diagrams to the estimated life over a given mission-in half a day. Huch more needs to be done in the area of preliminary design methods because the advancing stateof-the-art has made the design problem more complex and more aspects need to be considered in the preliminary phase. Cooling system design is an example. Turbine cooling is expensive. The coolant is lost to the cycle when it is extracted from a compressor stage and is discharged from the turbine. The discharge interferes with the mainstream flow. The need to provide flow paths through the engine and into the blades and vanes increases the complexity and cost of the machine. The designer naturally wants to use the smallest quantity of coolant possible and he would like to define that quantity as early as possible in the design cycle in order, for example, to be sure that the airfoils have sufficient internal flow area to pass the coolant. From an aerodynamic loading viewpoint, the turbine airfoils need to be slender. From a low heat transfer viewpoint, the opposite is desirable. Unfortunately, the current state-of-the-art puts this kind of caluclation out of the area of preliminary analysis. There is no fast and yet accurate method of determining the heat transfer to airfoils and endwalls and the available test data are far from comprehensive. Flow passages inside the airfoils are so complex that it is almost impossible to calculate the heat transfer coefficients. When the coolant is finally discharged, either as a film or through the trailing edge of a blade, the current state of knowledge does not permit an accurate prediction as to how it interacts with the free stream flow. The designer, therefore, must make do with a very preliminary estimate, with a view to improve it in the detailed design phase. Detailed Design The need for detailed design calculations has been recognized from the very early days of the gas turbine development. The turbine flow field is, after all, three-dimensional, with heat rransfer, mass addition, thick boundary layers, and transonic velocities; whereas the preliminary design calculations are essentially one-dimensional. Before the early 1950's, detailed design consisted of applying empirical correctj_ons to the preliminary design. The complexity of the flow field and the Navier Stokes equations precluded any analysis in greater depth. Closed form solutions were not (and still are not) apparent and the sheer magnitude of the calculations required put any numerical solution out of reach of the desk top calculators then availabl. In the early 19SO's, two events significant to gas turbine design occurred (2). Wu published his theory of S1-S2 stream surfaces, on which the flow is two-dimensional, and the second event was the emergence of the digital computer. The S1-S2 concept led to the development of radial equilibrium and streamline curvature analyses in the meridional plan for which the digital computers could generate solutions in a few hours. By the early 1960's, streamline curvature methods had been highly developed and were in everyday use, as indeed they are today, and computers could generate solutions in a few minutes. Turbine and compressor designers could economically obtain a much more accurate picture of the flow field which relied less extensively on empirical corrections. Still streamline curvature methods are limited in accuracy because they attempt to solve a three-dimensional problem by a series of two-dimensional approximations. Therefore, the need for a fully three-dimensional Navier-Stokes solution continued to be pressing, a\miting only a sufficiently powerful computer and the necessary skill to develop a tractable algorithm. Numerous investigators attempted solutions and by the early 1970's reasonably accurate solutions of turbomachinery flow fields were obtained for incompressible viscous flows with no heat transfer. One of the most successful is described in Reference 5. Figures 1 and 2 show some typical predictions of this program. Figure 1 is a contour plot of the kinetic energy loss at the exit plane of a turbine stator and Figure 2 is a plot of the "mixed out" (average bladeto-blade) efficiency at the same plane as compared with test data from a half inch downstream. The comparison can be seen to be quite good. In addition to aerodynamic performance, the program predicts heat transfer The case shown required about 40 minutes of CDC 6600 time and cost about $1, 000. This may seem expensive until one considers the cost of a test project to determine the equivalent aerodynamic and heat transfer information. One can readily see that three-dimensional fluid flow computations are now 2

4 TIP <l &ign use and validation of prediction methods, the AFAPL in partnership with the NASA Lewis Research Center is sponsoring a program to measure heat transfer to the endwalls of typical turbine cascades. Measurements will be taken at a variety of operating conditions, with and without coolant flow. Steady state heat transfer tests are expensive and time-consuming and it is very difficult to simulate engine conditions in a test rig, much less keep instrumentation intact at typical engine temperatures. Figure 1. "'.. HUB Exit Plane Survey \ I l: Dunn & Stoddard (3) demonstrated the capability of obtaining detailed heat transfer measurements to a turbine stator at actual engine conditions by using a shock tube as the hot gas source. They further showed the potential capability of obtaining aerodynamic measurements with the same apparatus. There is, however, some uncertainty with regard to the boundary layer fluctuation or intermittency phenomenom which may not be present in the short duration shock tube environment. They are currently engaged in applying the same technique to a complete stage. A fast, inexpensive technique such as this allows rapid comparisons of alternative designs, and because of the detailed heat transfer map it provides, allows the designer to put the coolant flow where it is needed without risk of over- or under-cooling. I> Figure 2...,_.; "',_ "b.oo 0.20 o. n o.so RELATIVE SPAN "Mixed-out" Efficiency 0.80 l.00 economically feasible as well as technically necessary, as is evidenced by its rapid adoption by the turbine engine industry. Beyond dispute, three-dimensional fluid flow computations are here to stay and extensive further growth is expected. Rapid advances at least through 1985, are projected with some serious obstacles yet to be tackled. For example, the boundary layers in turbomachinery are not well understood for the twodimensional case and much less so in three-dimensional cases. We have only begun to explore the phenomenon of rotor-stator interaction, which profoundly influences both the free stream and the boundary layer flow thereby, the heat transfer), and we will have to (and learn to account for the effect of discharged coolant on both flows. Naturally, we will want extensive experimental verification of all analytical methods. APPROACH The Air Force Aero Propulsion Laboratory the NASA, the Army, the Navy, other Government agencies, (AFAPL), and industry are conducting numerous research projects, both experimental and analytical, directed toward improving gas turbine component efficiencies. Some typical AFAPL programs are described below. In response to the need to minimize coolant flow and provide aerodynamic and heat transfer data for Research on airfoil boundary layers and cooling methods has provided the first photographic evidence of Taylor-Goertler vortices on the pressure surface of a turbine airfoil, and is yielding valuable insights on the transition from laminar to turbulent boundary layer flow. Another program has shown for the first time the trajectories of coolant flows discharged from a turbine rotor. Turbine aerodynamic and analytical research led to the development of the first viable 3-D viscous flow analysis, and to the LART stator, the first to be designed using a three-dimensional flow analysis. high pressure stage made up of this stator and a conventional rotor developed 92% efficiency at the design point, which vividly demonstrates the benefits to be obtained from such a design tool (4, 5). The current program seeks to develop a three-dimensional analysis in rotary coordinates, using this analysis, it with the existing stator. program, to design a rotor and to build the rotor and test In other phases of the tip leakage flow fields and the transient region between rotor and stator will be measured and analytical models developed. Other analytical work includes the Turbine Design System (6), a set of preliminary design analyses designed to be used through an interactive graphics terminal. A schematic is shown in Figure 3. Such a system increases designer productivity markedly because it eliminates the labor of punching data cards and it eliminates the wait for job turnaround. A Once the preliminary design is satisfactory, the user can produce an input set for both the 3-D flow analysis and the 3-D structural analysis. Additions to produce tapes for numerically controlled machine tools are feasible and are coming into use in some industries (7). OUTLOOK AND CONCLUSIONS Turbine design is entering a period of rapid and profound change caused by the simultaneous arrival of interactive graphics methods and a viable threedimensional flow analysis. By 1985, the typical preliminary design and analysis system will resemble that 3

5 Interactive graphics techniques.and viable three-dimensional methods in both flufd flow and structural analyses will together greatly increase the designer's productivity and reduce the need for development testing of new designs. This means that component efficiencies can be achieved at significantly reduced time and cost. REFERENCES 1. Zucrow, M. J., "Aircraft and Missile Propulsion," Vol. II, New York, John Wiley & Sons, Serovy, G. K., ed., "Fluid Dynamics of Turbomachinery," Course notes for ASME short course Fluid Dynamics of Turbomachinery, July 1973, Ames, Iowa. 3. Dunn, M. G. and Stoddard, F. J., "Measurement of Heat-Transfer Rate to a Gas Turbine Stator," ASHE 78-GT-119, Figure 3. Turbine Design System Schematic shown in Figure 3, with extensive use of graphics along the way. The detailed analysis will be fully threedimensional and account for rotation, heat transfer, mass addition, leakage flows, rotor-stator interactions, boundary layer transition, etc. It will be linked to the preliminary system as shown in Figure 4 and will be in daily use as a partial substitute for rig tests. Given a sufficiently robust computer, it may be possible to predict the flow through, and performance of, a complete stage. 4. Tall, W. A., "Understanding Secondary Flows," Secondary Flows in Turbomachines, AGARD Conference Proceedings No. 214, 1977, pp through Dodge, P. R. "3-D Heat Transfer Analysis Program," AFAPL-TR-77-64, 1 Oct 77, Air Force Aero Propulsion Laboratory, Wright-Patterson Air Force Base, Ohio. 6. Wysong, R. R., et al, "Turbine Design System," AFAPL-TR-78-92, Sept 78, Air Force Aero Propulsion Laboratory, Wright-Patterson Air Force Base, Ohio. 7. Roberts, D. I., "3-D Interactive Graphics Assist Diesel Engine Design," Mechanical Engineering, Vol. 100, No. 3, March 1978, pp Goldman, L. J. and McLallin, K. L., "Cold-Air Annular-Cascade Investigation of Core-Engine Cooled Turbine Vanes," NASA TH X-3224, April 1975, NASA Lewis Research Center, Cleveland, Ohio. 3-D Aero Analysis 3-D stress Analysis Figure 4. Advanced Turbine Analysis System 4