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Highly Loaded Axial-Flow Compressors - History And Current Development

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Highly Loaded Axial-Flow Compressors - History And Current Development A. J. Wennerstrom Aero Propulsion and Power Laboratory, Wright Research and Development Center (AFSC), Wright-Patterson Air Force Base, OH 45433-6563 Highly Loaded Axial Flow Compressors: History and Current Developments This paper discusses appro...

Highly Loaded Axial-Flow Compressors - History And Current Development
A. J. Wennerstrom Aero Propulsion and Power Laboratory, Wright Research and Development Center (AFSC), Wright-Patterson Air Force Base, OH 45433-6563 Highly Loaded Axial Flow Compressors: History and Current Developments This paper discusses approaches taken over many years to achieve very high loading levels in axial-flow compressors. These efforts have been associated predominantly with aircraft turbine engines. The objective has been to reduce the size and weight of the powerplant, to increase its simplicity and ruggedness, and, whenever possible, to reduce cost. In the introduction, some fundamentals are reviewed that indicate that increased work per stage can only be obtained at a cost of increased Mach number, increased diffusion, or both. The earliest examples cited are some ambitious development programs of the 1950s and 1960s. Some innovative schemes to increase diffusion limits are described that took place in the 1960s and 1970s. Major ad- vancements in dealing with higher Mach number were made in the 1980s. Finally, a few thoughts directed toward potential future developments are presented. Introduction One of the most obvious ways to reduce the weight of the axial-flow compressor in an aircraft turbine engine is to reduce the number of stages it requires. Inasmuch as the engine cycle will dictate the overall pressure ratio required, a reduction in the number of stages inevitably means that each stage must do more work. Maximizing the work per stage while retaining ac- ceptable overall performance has been a goal of the aircraft engine designer ever since an axial-flow compressor was first incorporated into an aircraft engine. In this section, I want to illustrate what more work per stage means to the compressor. Then in subsequent sections I will illustrate several research and development efforts to which this has led and then will conclude with a few comments concerning future prospects. The most fundamental expression for work per stage is the Euler Equation of Turbomachinery, which, neglecting radius change, can be written for compressors Ah, = mva-vn) (1) This tells us that the work per stage is directly proportional to the wheel speed and the change in absolute swirl velocity across the rotor. The upper limit for wheel sped is most often defined by some structural limitation. However, since as wheel speed increases, relative flow velocities and hence Mach numbers increase, in some instances such as high bypass tur- bofans, wheel speed may be Mach number limited for reasons of thermodynamic efficiency . The permissible change in swirl velocity across a rotor is, on the other hand, almost always limited by some aerodynamic constraint. This constraint may be Mach number or some loading parameter related to boundary layer separation. Contributed by the International Gas Turbine Institute and presented as an invited lecture at the Ninth International Symposium on Air Breathing Engines, Athens, Greece, September 3-8, 1989. Manuscript received at ASME Head- quarters February 5, 1990. Losses, and hence efficiency, and also stall margin are also related to work per stage. Although losses result from many different phenomena, the majority of losses can be divided in- to those related to Mach number (shock waves) and viscous losses related to diffusion acting on boundary layers. Some of the earliest indices of loading considered only inlet to exit con- ditions across a blade row. These guidelines were arrived at by examination of a very limited range of low-speed data. They included the static pressure rise coefficient Pi~P\ -<0.6 PW\ and the de Haller number W1 W, >0.72 (2) (3) The former defined the maximum practical static pressure rise as a function of inlet dynamic head. The latter defined a minimum exit-to-inlet relative velocity ratio to achieve ac- ceptable performance. Neither of these indices were correlated against losses. The most useful diffusion parameter to be in- troduced was the diffusion factor presented by Lieblein et al. (1953). This not only provided some guidance relative to prac- tical upper limits for conventional blade rows, but it also proved a useful correlation parameter for relative total pressure viscous losses. The diffusion factor, again neglecting radius change, can be written D=\ W2 Vn-Vn 2a W, (4) Through a series of simplifying assumptions in its derivation, Journal of Turbomachinery OCTOBER 1990, Vol. 112/567 Copyright © 1990 by ASME Downloaded 17 Feb 2012 to 159.226.48.114. Redistribution subject to ASME license or copyright; see http://www.asme.org/terms/Terms_Use.cfm the above factor is intended to be an approximate representa- tion of the local suction surface diffusion factor defined by ^\nra] ~~ ~ W •W, w rr avg It is interesting to compare these three parameters for the ex- treme case of an impulse blade row. For the sake of simplicity, let us further consider the flow to be incompressible and to have no inlet absolute velocity swirl, i.e., Vei = 0. Since im- pulse blading has no static pressure change, the pressure rise coefficient P2-P1 -^-pw\ An impulse blade row in incompressible flow will have sym- metric inlet and exit velocity diagrams in the relative frame. ri and the de Haller number W2 W, = 1.0 Therefore, according to these criteria, the blade row has no loading whatsoever. However, the diffusion factor can have a substantial value and will be D- V- U 2a W, aW, or, in terms of angles, D = sin /3, Although the magnitude of this diffusion factor may be mean- ingless because the existing correlations were derived for such different blading, the diffusion factor does correctly indicate that there can be a large amount of diffusion even in an im- pulse blade row. The fluid will accelerate to a peak velocity somewhere on the suction surface while turning and then must inevitably diffuse to a much lower value to achieve equilibrium exit conditions. In order to see the combined effect of wheel speed and dif- fusion on stage work capacity, we can combine the Euler equation, equation (1), with the equation for diffusion factor, equation (4). This leads to W-, Mil = 2aUWAD-\+-~) We see from this that work capacity is directly proportional to wheel speed, is nearly proportional to inlet relative velocity, and varies in some proportion with diffusion factor; if W2/W1 = 1.0 it would be directly proportional. One also observes that stage work capacity is directly proportional to solidity. However, this is subject to several practical con- straints in that skin friction losses will increase with wetted surface, increased blade blockage will reduce choking mass flow, and finally the weight will increase in more than direct proportion to the blade count. Hence, it is readily apparent why all efforts to increase stage work capacity have concen- trated on increasing wheel speed, Mach number (relative velocity), and diffusion. Supersonic Compressors Early History. As was pointed out in the previous section, the most obvious way to increase work per stage was to in- crease wheel speeds, whirl velocities, and hence Mach numbers. This was pointed out many years ago and the idea of operating a compressor at supersonic relative velocities with normal shock waves in the blading is generally first credited to Weise (1937) in Germany. Two excellent summaries of early work with supersonic compressors are presented by Klapproth (1961) and Erwin (1964). Klapproth made the astute observation that the early design approach for supersonic compressors was developed more from supersonic diffuser design criteria than from conven- tional compressor design criteria. As a consequence, early work with supersonic compressors was not a logical evolution of work with lower speed machines but rather attempted to make a quantum leap forward with little regard for past tur- bomachinery experience. Hindsight has shown that many of the loading criteria developed for lower-speed machinery have retained a remarkable validity for high-speed machinery, as il- logical as it might seem. The more successful machines built over the years have combined concepts of dealing with higher Mach numbers with older, more conventional concepts of blade aerodynamic loading limits. Most work concerning supersonic compressors has already been published in the open literature. What I would like to present here, in approximately chronological order, are four early, very serious development efforts that have not previous- ly been published because they were originally classified. Each of these incorporated a supersonic compressor, which went through extensive engineering development. All four represented extremely ambitious exploratory or advanced development programs, had varying degrees of success because of that, and hence never found their way into any pro- duction engine. The J-55 Turbojet. This is probably the first gas turbine to run as an engine incorporating a supersonic compressor. Its development was initiated in 1947 with the objective of power- ing a target drone. Its compression system was unique in in- corporating a single-stage supersonic axial compressor of the shock-in-rotor type, designed to provide a 2.75 total pressure ratio at a corrected tip speed of about 488 m/s (1600 ft/s). It included inlet guide vanes, which provided some counterswirl at the hub and incorporated a tandem-bladed stator. The overall engine never achieved enough of its design goals to be put into production. However, by the conclusion of the con- tract, the compressor had achieved a total pressure ratio of ap- proximately 2.9 at an isentropic efficiency of about 0.76 at design speed. Lower in its speed range its efficiency peaked at about 0.82. One of the major development prob- lems was achieving adequate performance from the supersonic compressor. In 1947, the contractor, Fredrick Flader, Inc., was optimistic enough to state, "we are convinced that this development has progressed to a point where its incorporation Nomenclature D = diffusion factor Aocai = local diffusion factor h = enthalpy p = static pressure U = wheel speed Ve = absolute swirl velocity W = relative velocity W W P = peak suction surface relative velocity average suction surface relative velocity relative flow angle measured from axial static density a = solidity (chord/spac Subscripts 1 = blade-row inlet 2 = blade-row exit t = stagnation quantity 568 / Vol. 112, OCTOBER 1990 Transactions of the ASME Downloaded 17 Feb 2012 to 159.226.48.114. Redistribution subject to ASME license or copyright; see http://www.asme.org/terms/Terms_Use.cfm Table 1 Counterrotating compressor design characteristics at sea-level takeoff conditions Low-pressure rotor Design tip speed Total pressure ratio (design) Adiabatic efficiency Inlet diameter ratio Exit diameter ratio High-pressure rotor Design tip speed Total pressure ratio (design) Adiabatic efficiency Inlet diameter ratio Exit diameter ratio Overall compressor Overall pressure ratio (excluding inlet losses) Overall adiabatic efficiency Compressor o.d. Compressor weight flow (design) Weight flow/frontal area Impulse type 442 m/s (1450 ft/s) 3.69/1 0.847 0.65 0.772 Shock-in-rotor type 266 m/s (874 ft/s) 1.829/1 0.623 0.772 0.827 Statorless 6.75 0.775 406 mm (16 in.) 17.7 kg/s (39.0 lb/s) 135.7 kg/s/m2 (27.8 lb/s/ft2) Fig. 1 Near-midspan cross section of rotor blading in this design proposal is w a r r a n t e d . " History showed this conclusion to be somewhat premature. These bits of information were derived from Downs (1953) and Blanton (1989). Curtiss-Wright Counterrotating Compressor. This was probably one of the most ambitious of all supersonic com- pressor designs ever attempted. The effort took place in the mid-1950s and the design was reported by Sabel and Sabatiuk (1957a). It was designed for a flight Mach number of 3.0, to accept an axially supersonic inlet Mach number at this condi- tion utilizing an integral fixed-geometry free-stream inlet, and to be able to take off from a sea-level static condition with a choked inlet. One of its major goals was to eliminate the necessity for a heavy, complex, variable-geometry inlet and its controls and actuating mechanisims. Its objective was to employ counterrotating rotors to avoid the necessity for any inlet guide vanes or stators and to achieve an adequate pressure ratio over its flight regime through this means. Its projected design features are summarized in Table 1 for the sea-level takeoff condition. Most notable is the fact that it was proposed to do this with an impulse first rotor and a shock-in-rotor second rotor. As with several of its predecessors, supersonic diffuser design practice guided design of the second rotor. Conventional compressor loading criteria were completely ignored. Because of the sea-level static takeoff requirement, the leading edges of the first rotor were staggered at an angle of two or three degrees at this con- dition. However, at the flight cruise condition, the first rotor was anticipated to operate at negative incidence and with a swallowed shock pattern. The first rotor incorporated all of the technology con- ceivable at the time. This included leading edge sweep de- signed to keep the sonic line coincident with the surface of revolution defined by the leading edge. An axisymmetric stream filament design technique was used employing simple radial equilibrium theory. As understood today, this would exclude internal computing stations and streamline curvature • effects. The choice of solidity and aspect ratio was based on previous recent NACA experience. The mean solidity of the first rotor was approximately 2.8. Maximum thickness varied from 3.3 percent chord at the tip to 7.9 percent chord at the hub. A method-of-characteristics computation was made to insure a reasonable probability of passing design flow. The second rotor was designed for an average relative inlet Mach number of 2.34. The internal one-dimensional area con- Fig. 2 Photograph of high-pressure rotor traction of the blading was designed to permit starting a relative approach flow of Mach 2.0. Twenty-five degrees of leading-edge sweep was incorporated to enhance starting of the supersonic flow. Today we would say that this rotor was designed with significant precompression, and from the throat to the trailing edge plane an equivalent cone angle of approx- imately eight degrees was employed. The mean solidity was approximately 2.05. The same aerodynamic design technique was employed as for the first stage. An approximately midradius cross section of both rotors is shown in Fig. 1. The second rotor is shown assembled in Fig. 2. Interestingly enough, diffusion factors were calculated. At sea-level take-off conditions, they were 0.258 for the low- pressure rotor and 1.00 for the high-pressure rotor. For the first rotor of impulse design, the diffusion factor is not ter- ribly meaningful. However, the diffusion factor of 1.00 for the second rotor is typical of so many unsuccessful designs of that era. When one looks back on the overall experience, it is remarkable to what degree performance does correlate with diffusion factor under circumstances far removed from those under which it was derived. The experimental results of sea-level static testing were reported by Sabel and Sabatiuk (1957b). Choking in the first rotor limited weight flow to approximately 79 percent of the design value. Measured performance recorded with both spools at design speed was an overall total pressure ratio of 3.24 and an isentropic efficiency based on temperature rise of 73.3 percent. This efficiency seems surprisingly high when it is noted that the total pressure ratio is only 48 percent of the design value. Observation of casing static pressures presented in Fig. 3 shows that most of the static pressure diffusion ac- tually occurred downstream of the second rotor rather than within it. Inasmuch as the exit plane instrumentation was within this region, it is likely than when exit mixing losses are taken into account, the true efficiency would be much lower. The casing static pressures tell another interesting story. The static pressure ratio across the first impulse rotor was intended to be about unity and this was the result achieved. However, the static pressure ratio intended to exist across the second rotor was about 11.7 according to the design. In fact, the ex- perimental results show a negligible static pressure rise also Journal of Turbomachinery OCTOBER 1990, Vol. 112/569 Downloaded 17 Feb 2012 to 159.226.48.114. Redistribution subject to ASME license or copyright; see http://www.asme.org/terms/Terms_Use.cfm Fig. 3 Casing static pressure distribution Fig. 5 Close-up of assembled high-pressure rotor viewed along span Fig. 6 TS-10 supersonic rotor Fig. 4 DLD gas generator supersonic compressor blade across the second rotor. From inlet to exit, an overall static pressure ratio of only about 2.0 was achieved and all of this occurred over a distance of several annulus widths downstream of the second rotor. Several additional tests were made with this rig. However, a major facility-related mechanical failure destroyed the com- pressor and this program was terminated without further significant accomplishment. Direct Lift Demonstrator. This was a program ac- complished by the General Electric Company in the mid-1960s. Its objective was to develop the major components of a turbofan engine to a level of technology consistent with direct lift (VTOL) requirements and then to assemble them as an engine and test it in a typical environment. It was a two- spool turbofan. The low-pressure spool had a single transonic fan stage. The high-pressure spool had a single supersonic compressor stage. My remarks will be limited to aerodynamic aspects of the supersonic core stage. The data used were presented by Conliffe (1971). The supersonic stage was a shock-in-rotor/shock-in-stator configuration. Its design characteristics included a design total pressure ratio of 3.15 at a corrected tip speed of 420 m/s (1378 ft/s) and with a projected isentropic efficiency of 75 percent. The inlet hub/tip radius ratio was approximately 0.82. Relative Mach numbers ranged from about 1.2 to 1.4 across the leading edge of the rotor and from about 1.4 to 1.5 across the leading edge of the stator. Design static pressure ratios at the outer casing for both rotor and stator were approximately 1.9. Rotor solidity was approximately 2.2; stator solidity was about 4.8. Mean aspect ratios were approximately 0.88 and 0.27 respectively. An example of the rotor blading is shown in Fig. 4. A close-up of the assembled rotor is shown in Fig. 5. The performance achieved for the best configuration subse- quently used in the gas generator was a total pressure ratio of about 3.5 at 71.4 percent isentropic efficiency. A large number of configurations were examined during the development pro- gram. The majority were aimed at improving the stator per- formance. Although the indicated performance of the rotor had been quite good throughout the program, the stator total pressure recovery and discharge flow distribution remained quite poor. In this instance, the designers observed generally accepted limits for loading parameters such as diffusion fac- tor. However, once again, the inability of a stator to accept Mach numbers and static pressure ratios as high as those of rotors had been demonstrated and nothing the designers at- tempted altered this situation substantially. Turboaccelerator (TS-10). This was a progr
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