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AERODYNAMIC DESIGNDESIGN AND TESTING OF AN AXIAL FLOW Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 WITH PRESSURE RATIO OF 23.3:1 FOR THE LM2500+ GAS ' 1111111 111111 A.R. Wadia, D. P. Wolf and F. G. Haaser GE Aircraft Engines Cincinnati, Ohio 45215

ABSTRACT (1998). The LM2500 , derived from the CF6-61TF39 The LM2500+ gas turbine, rated between 39,000 to 40,200 aircraft engines, has also leveraged off technology shaft horsepower (shp), was introduced for field service in 1998. development to facilitate the increase in its industrial power rating This growth aero-derivative gas turbine is suitable for a variety of from the original 24,000 shp to the current 31,200 shp. Market power generation applications, such as co-generation and com- studies initiated in the late eighties and early nineties showed that bined cycle, as well as mechanical drive applications. At the heart the LM2500 industrial gas turbine needed additional power of the LM2500+ 25% power increase is an up-rated derivative 17- (39,000 shp at ISO conditions) to meet customer requirements stage axial compressor. This paper describes the aerodynamic (Farmer, 1994). This up-rated power version of the LM2500 was design and development of this high pressure ratio single spool named the LM2500+ gas turbine. The LM2500+ gas turbine 3D compressor for the LM2500+ gas turbine. The compressor is cutaway presented in Figure 1 highlights the key modifications derived by zero-staging the highly efficient and reliable LM2500 made to the engine relative to the LM2500 base engine. compressor to increase the flow by 23% at a pressure ratio of 23.3:1. The aerodynamic efficiency of the compressor is further improved by using three-dimensional, custom-tailored designs similar to those used in the CF6-80C2 high pressure com- pressor. The compressor achieved a peak polytropic efficiency above 91 percent, meeting all its operability objectives. The tech- nical requirements and overall aerod ynamic design features of the compressor are presented first. Next, the zero stage match point selection is described and the procedure used to set up the vector diagrams using a through-flow code with secondary flow and mixing is outlined. Detailed design results for the new transonic in the compressor using three-dimensional viscous analy- sis are presented. The compressor instrumentation and perfor- mance test results are discussed. The performance of the zero Figure 1: LM2500+ gas turbine unique features. stage is separated from that of the baseline compressor with the CF6-80C2 airfoils to show the improvement in efficiency with the In March 1994, after a series of preliminary design studies on new airfoils. how to achieve the required power increase, it was decided to launch the LM2500+ which was chosen from four candidate con- NOMENCLATURE figurations based on a cost-and-risk assessment comparison. The Blisk = Bladed Disk preliminary design team evaluated power enhancement tech- C = Airfoil Section Chord niques such as inter-cooling, inlet supercharging, recuperation and IMM = Radial Immersion (0=tip, 1=hub) other refinements. The team, however, decided on the basis of Tmax = Airfoil Section Maximum Thickness design simplicity, program schedule, technology risks and devel- Z = Axial Distance (inches) opment cost and customer price that increasing the inlet mass flow through the engine was the simplest and most conservative INTRODUCTION way of increasing the power output. The increased mass flow Application of aero-engine technology to ground-based gas could be achieved by zero-staging the current production LM2500 has increased rapidl y, especially in the last decade, as compressor. Simultaneously, the increase in the turbine rotor inlet documented in the works of Scalzo (1988), ICashiwabara (1990), temperatures could be minimized to approximately 35 degrees C Sehra (1991), Smed (1991), Janssen (1995), and Stringham (65 degrees F), (Valenti, 1998) by going to a more efficient com-

Presented at the International Gas Turbine & Aeroenglne Congress & Exhibition Indianapolis, Indiana — June 7—June 10, 1999 pressor design using three-dimensional analytical tools and incor- COMPRESSOR AERODYNAMIC DESIGN FEATURES porating custom-tailored compressor airfoils from the CF6-80C2 The detailed aerodynamic design of the original aircraft engine. It was apparent that this design approach best met CF6-61LM2500 compressor has been reported by Klapproth. the goals of using proven technology at minimal risk. Also, Miller, and Parker (1979). To achieve the increased power output keeping a strong fundamental LM2500 design heritage facilitated rating, the LM2500+ required a 23% increase in airflow. The 23% product support while meeting the objectives of a base load hot increase in flow was achieved by adding an additional compres- section inspection interval of 25,000 hours with engine overhauls sion stage (zero stage) to the LM2500 and flaring the flowpath in at 50,000 hours. front of the existing compressor to form an overall 17-stage axial compressor unit. Figure 2 shows the changes incorporated into the Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 COMPRESSOR TECHNICAL REQUIREMENTS LM2500+ relative to the base compressor. As a result of zero- SUMMARY staging of the compressor, the LM2500+ increased in length by The performance requirements for the LM2500+ compressor 34.3 cm (13.5 inches) and the weight of the engine increased by were less stringent relative to those required by commercial or about 363 kg (800 lbs). Table 1 shows the comparison of key aero- military aircraft engines. While aircraft engines have multiple dynamic design parameters for the LM2500 compressor with operating points such as take-off, cruise, etc., where performance those selected for the design of the LM2500+ compressor. is crucial, the LM2500+ is required to operate at close to its peak 10.. 'MY. *ran, Anus*. 0117eva TAM efficiency near its high-speed design point. While no specific per- labdwplint Nepal Nth Biz& Bird Ill. Now MSS MY Chat formance requirements at other speeds were specified, it was Slop 0 ELIStShislt desirable for the compressor to preserve good efficiency over a ExIntIsd I,,c Ca." SS Ple Stow I0-13 Spool' Own range of speeds. ta, lisakle Pad Saga favnglianme Additionally, the LM2500+ compressor was also required to Nan 11-15 Spool operate stall-free with both a Single Annular (SAC), which results in a smooth compressor operating line, and with a /111111110110 Dry Low Emissions (DLE) Combustor, which results in a com- alliratrar11111 "s"" pressor operating line with steps corresponding to the staging in

lbw 0 the combustor. The compressor operating line can vary by as SONO Wel Alnluel MI COP Lel Ws tow I I/ 2 Eftlo much as 2 percent below and above the nominal operating line, -1115 In. long. Stage 1.11 CFIMOC2 Skits Ver. Maps 201 MAO= 131.1s. from the start of the combustor staging to the end of the staging for Irryned Ucy sequence, respectively. Figure 2: LM2500+ high pressure compressor improvements The customer-supplied inlet systems used with industrial gas relative to the base (LM2500) compressor. turbines, such as the LM2500+, are generally quite aerodynami- cally "clean" and use a light wire screen mesh to prevent any large Table I: Compressor Aerodynamic Design Operating Point objects being ingested by the compressor. Maneuvers and cross- Parameters LIME LAMLE .Dagi wind inlet distortion issues are almost non-existent on these Shaft Horsepower 31,200 39,000 25% land/marine-based engines, thus easing the operability require- Inlet Corrected Flow, kg/s 68 (150 lb/s) 84.5 (186 Ibis) 23% ments. Installation design manuals suggest the inlet distortion 9,586 1.5% index to be of the order of 2 percent or less as most of the opera- Inlet Corrected Speed (rpm) 9,418 tion is with a straight bellmouth or radial volute. To account for Pressure Ratio 18.8 23.3 23% any inlet distortion that might be encountered in the field, Polytropic Efficiency 88.9% 91% 2.36% the compressor was designed with a slight tip radial (i.e., total pressure deficit at the tip) inlet total pressure profile to realistical- ZERO STAGE MATCH POINT SELECTION ly simulate the tip aerodynamic loading level on the zero Technical information on zero-staging in the stage blade. open literature is limited. Some of the principles in the develop- Acoustics plays an economic role in land/marine-based ment of front stages of axial flow compressors has been reported systems design, and the goal for the LM2500+ was to maintain the by Eisenberg (1993) and Katoh (1993). A recent compressor zero- same inlet noise sound pressure level in spite of the 23% higher staging application to the Taurus 60 axial flow compressor that airflow. This requirement set the vane/blade ratios and the axial increases the inlet mass flow by approximately 20% and raises spacing between the rotor and the stator using an "acoustic cut- the pressure ratio from 11.2:1 to 16:1 has been reported by off" design criteria. Van Leuven (1994) and Rocha, Saadatmand, and Bolander (1995). The compressor operating line was set with a minimum of 12 The CF6-6/LM2500 compressor performance map has been percent stall margin to account for any operating line migration presented in the paper by Klapproth, Miller, and Parker (1979). that would Deem' in service during the life of the engine. Start The LM2500 compressor achieved a peak polytropic efficiency of times for the engine are of the order of two minutes, which is less 90.7% along the engine operating line at a compressor inlet cor- stringent than in aircraft engines. A four-degree open stator stall rected flow of 60 kg/s (132 lbs/sec) and pressure ratio of 15:1. The margin requirement was conservatively set for the LM2500+ to 88.9 percent polytropic efficiency shown in Table 1 (does not account for Variable Stator Vanes (VSV) control/rigging variation include the compressor exit diffuser losses) represents the com- and deterioration in the field. pressor performance at the maximum power operating point of the

2 LM2500 engine, which is different from the peak efficiency point airfoil camber and stagger angles for all the LM2500 blading has on the . been published by Klapproth et. al. (1979). Table II summarizes To achieve maximum compressor performance with the zero some of the key airfoil geometry and vector diagram quantities for stage, the point corresponding to the peak efficiency of the exist- the first three stages of the LM2500+ compressor. ing LM2500 compressor would be selected as the match point for - . Table II: Rotor and Stator Pitch Line the LM 2500+ compressor design. This sets the zero stage pressure Design Parameters Summary_ ratio such that the LM2500+ compressor achieves its overall com- Parameter TA Rotor Q $tator Q Boma pression ratio goal. In order to achieve this optimum pressure $tator I. Rotor 2 ratio, the zero stage's "stage effectivity," as calculated using the Solidity 1.045 1.582 1.281 1.176 0.974 0.915 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 approach formulated by Koch (1981), exceeded that demonstrat- Aspect Ratio 6.04 1.390 4.000 2.337' 3.708 2.609 ed by the first stage of the compressor in GE Aircraft Engine's Inlet Mach No. 0.489 1.018 0.706 0.826 0.727 0.771 Energy Efficient Engine (E3), which has a similar overall pressure Diffusion Factor 0.423 0.267 0.348 0.353 0.355 ratio as that for the LM2500+. • The term "stage effectivity" is a correlation that is used inter- A preliminary design analysis resulted in the selection of 16 nally within GE Aircraft Engines to assess a stage's loading rela- blades (Rotor 0) for the zero stage. The selection of the number of tive to the loading at stall. It is synonymous to the peak stage inlet guide vanes and the number of zero stage vanes (vane/blade static pressure rise coefficient and has been correlated versus a cut-off ratio), and the axial gaps between the inlet guide vanes, parameter involving stage average values of solidity and aspect Rotor 0 and Stator 0 were set by the technical requirements for an ratio. The correlation serves as an aerodynamic loading limit, and "acoustic cut-off" in order to prevent higher inlet noise levels. accounts for factors such as blade speed, axial velocity, reaction This resulted in larger-than-normal axial gaps on both sides of the ratio, clearance and Reynolds number to form a systematic zero stage blade, as shown in Figure 2. These large gaps are detri- method for analysis or prediction of performance. mental to compressor performance, making the aerodynamic The selection of the pressure ratio, for the zero stage, was an design of the compressor even more challenging. iterative process to achieve a balance between the design point The following sections summarize the detailed design of the pressure ratio and its "stage effectivity" at stall. This iterative new front stages of the LM2500+ compressOr with emphasis on process yielded a zero stage pressure ratio of 1.438, which is the zero stage and the first stage blade designs : Three-dimension- slightly less than optimum. al viscous analysis using the computer program :developed by The efficiency potential of the zero stage was determined lennions and Turner (1993) was used extensively in the design of using the principles outlined by Koch and Smith (1975) and the the compressor blading. corresponding work input assigned to it. The resultant exit tem- perature and pressure from Rotor 0 provided the input to calculate FRONT FRAME AND INLET GUIDE VANE DESIGN the inlet conditions into the following stator. Inlet corrected flow The LM2500+ front frame is very similar to the original and speed at Stator 0 inlet were then calculated and compared to engine frame. It is a five-strut (4 thin struts and 1 thick strut), 17-4 the measured flow at speed along the base LM2500 compressor steel casting and retains the same inner and outer front forward operating line. The calculations converged quickly to provide a flange configuration for inlet commonality as shown in Figure 2. realistic axial match between the new zero stage and the down- The frame's inner and outer flowpath walls were flared to increase stream existing compressor. the flow area through the frame to prevent a higher The above-mentioned approach, substantiated by past expe- pressure loss due to a possible flow restriction. The original front rience on the E 3 engine and test data from the LM2500 compres- frame struts are bi-convex airfoils, which for the LM2500+ were sor, resulted in matching the LM2500+ zero stage to the original made more aerodynamic using NACA 65-series airfoils. LM2500 compressor at an inlet corrected flow of about 62.7 kg/s Thickness and airfoil contour changes lowered the front frame (137 lbs/sec) (corresponding to an inlet corrected speed of 8,950 losses as verified by three-dimensional viscous analysis of rpm) at 90.3% polytropic efficiency. the frame. The Inlet Guide Vane (IGV) airfoil design was done with a COMPRESSOR DESIGN VECTOR DIAGRAMS conventional profile. The correlation of NACA 63-series airfoil The vector diagrams used to design the new front stages of cascade data by Dunavant was used with the IGV solidity varying the LM2500+ compressor were derived from a data match of a from 0.89 at the tip to 1.2 at the hub. The IGV exit swirl was fully instrumented, CF6-80C2 core engine using the circumferen- varied almost linearly from about 18 degrees of pre-swirl at the tip tially averaged, through-flow code with secondary flow and to -15 degrees of counter-swirl at the hub. The tip pre-swirl pro- mixing (Adkins and Smith, 1981). Next, the new zero stage was vided relief to the downstream rotor at the tip, added to the through-flow analysis. The new set of vector dia- while the hub counter-swirl helped lower the hub Mach number grams obtained were such that the velocity triangle quantities — into the zero stage vane. The IGV design was important because such as the inlet relative flow angle, inlet relative Mach number it has to deliver the required swirl distribution to the zero stage and inlet meridional Mach number into stage 2 blade — were the blade at the design condition and it has to be able to operate ade- same as the data match values. quately at part speed when the airfoil is closed by as much as 60 The axial distribution of solidity, aspect ratio, axial velocity, degrees. The IGV airfoil was analyzed at the design point with the diffusion factor, Mach number, inlet and exit flow angles and three-dimensional viscous analysis. This analysis was done to

3 verify the magnitude and radial distribution of pressure loss The inlet relative Mach number is transonic over most of the assumed in the circumferentially averaged through-flow analysis blade span of the LM2500+ Rotor 0. The efficiency of a transon- and its ability to meet the exit swirl requirements. ic blade is heavily influenced by shock losses, which may exceed the losses due to cascade diffusion and secondary flow effects. ZERO STAGE BLADE DESIGN The Mach number just ahead of the leading edge passage shock Preliminary mechanical design studies showed that dovetail can be influenced by the shape of the blade suction surface ahead stresses on a traditional zero stage blade design would limit the of the shock. Increasing the average suction surface angle, as minimum radius ratio to 0.45. A blisk version of the rotor permit- measured from axial, ahead of the shock reduces the average ted a reduction in the radius ratio to 0.368. This reduction in Mach number upstream of the shock through external compres- radius ratio also provided some aerodynamic performance bene- sion and should reduce the shock losses. However, this type of Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 fits by lowering the inlet specific flow. The blisk also provided a airfoil can result in a reduced cascade throat area. If the throat is parts count reduction by replacing 40+ parts with a single part, too small, the cascade will not pass the design flow and may not and eliminated the wear issue of a midspan shroud, blade and disk achieve the attached shock pattern desired for minimum loss. The and dovetails. Although blisk technology has been in existence for lessons learned in transonic rotor performance studies by Wadia 20+ years, its inclusion in the LM2500+ represents the first intro- and Copenhaver (1996), with different cascade area ratios, were duction of a blisk to GE Aircraft Engines' Marine and Industrial applied to set the throat margin, internal contraction and trailing engines product line. edge camber of the zero stage blade. The blade was designed with Table III shows the comparison of the geometric and aero- a 5 percent throat margin. dynamic design parameters between the zero stage rotor for the The design was further analyzed with a three-dimensional LM2500+ compressor and the first stage blade from the GE viscous code to get more definitive results on the effect of custom Aircraft Engines E3 compressor. As illustrated by the thickness tailoring the mean camber lines and to verify that the cascade comparisons in Table HI, the LM2500+ blades were considerably would pass the design flow and deliver the design intent exit thicker (ruggedized) to increase Foreign Object Damage (FOD) radial profiles of total pressure and temperature. tolerance. The location of maximum thickness was also moved The flow computed by the three-dimensional analysis was forward on the LM2500+ airfoil, in a similar manner as illustrat- 1.05% higher than the design flow rate. This difference between ed in the paper by Wadia and Law (1993), to provide improved the design and calculated flow rate is consistent with that between resistance to leading edge foreign body impact. the three-dimensional analysis and measured flow on other similar transonic blade row designs. Table III: Zero Stage Blade Key Aerodynamic The radial profiles of total pressure, temperature and adia- Design Parameters batic efficiency (at Stator 0 leading edge) calculated by the three- LM2500+ GEAE E3 dimensional viscous analysis (3D) are illustrated in Figure 3. The Parameter Rotor 0 Rotor 1 total pressure profile is hub strong and agrees well with the design Number of Blades 16 28 intent also shown in Figure 3. The calculated efficiency was Inlet Corrected Flow (kg/s) 84.5 (186 lb/s) 54.5 (120 lb/s) slightly higher relative to the design vector diagrams. Inlet Specific Flow (lb/s-sq.in.) 36.0 36.2 Figure 4 shows the isentropic Mach number distribution on Stage Pressure Ratio 1.438 1.65 the pressure and suction surfaces of the LM2500+ zero stage Inlet Corrected Tip Speed (ft/s) 1,363 1,495 blade. Figure 5 shows the corresponding calculated shock struc- Inlet Tip Pre-Swirl (degrees) 18 14 ture at two immersions along the blade span. Near the tip, a two- Inlet Hub Counter-Swirl (degrees) -15 0 shock system was selected over a single leading edge shock Inlet Relative Tip Mach Number 1.19 1.32 pattern to obtain a balance between the efficiency and stability of the blade. It was recognized that the peak efficiency would occur Inlet Radius Ratio 0.368 0.52 on a slightly higher operating line where the passage shock Aspect Ratio 1.39 1.51 merged with the leading edge shock. At the midspan, Figure 5 Pitch Solidity 1.59 1.71 shows a single passage shock structure with the shock intersecting 0.46 Pitch Diffusion Factor 0.44 the pressure and suction surfaces at 30 and 65 percent axial dis- lip, Pitch, Hub Tmax/C .037, .088, .130 .024, .052, .096 tance from the leading edge of the airfoil. respectively. Figure 6 shows the isentropic blade surface Mach number Rotor 0 was designed to the vector diagrams created from a distribution near the hub of the blade. The hub scalloping provid- data match of the CF6-80C2 core compressor as reported earlier. ed the required relief to the 13 percent thick hub section by low- Transonic airfoil design principles presented in Wadia and Law ering the average Mach number and resulted in a shock-free (1993) were applied to custom tailor the mean camber lines to design by keeping the blade surface Mach numbers below unity. alleviate some of the performance penalties associated with the No flow separation at the hub was observed in the three-dimen- ruggedization. The detrimental effect on performance due to the sional calculations. large increase in thickness, especially near the hub, was also reduced by scalloping (area-ruling) the hub flowpath within the Figure 7 shows the zero stage blisk used in the first build of blade as shown in Figure 2. the LM2500+ engine.

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RADIAL PROFILE OF TOTAL PRESSURE PRIE3311Re SURFACE SUCTION SURFACE (PRETEST STATOR SCHEDULE) 9 S 7 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021

2 —0- Design Intent 1 • - 30 Design

0 16 17 18 19 20 21 22 23 24 Figure 4: LM2500+ Rotor 0 isentropic Mach number contours TOTAL PRESSURE (INLET-TO-STATOR 0 LE) (PS(A) on the pressure and suction surfaces. (a)

ISEMTROPIC MACH MOSER CONTOURS RADIAL PROFILE OF TOTAL TEMPERATURE (PRETEST STATOR SCHEDULE)

...... 4- - • 8 •

7 •

—0— Design Intent

1 - -.t- SODeign 0 575 580. 585 590 595 600 605 TOTAL TEMPERATURE (INLET-TO-STATOR 0 LE) (DEG-R) Figure 5: LM2500+ Rotor 0 blade passage shock structure at (b) 8 and 45 percent immersions.

RADIAL PROFILE OF ADIABATIC EFFICIENCY (PRETEST STATOR SCHEDULE) 1.2 9 IMM = 0.92 8 w1.0 7

2 1 Lui 0.2 0 04 0.5 06 0.7 0.8 0.9 10 ADIABATIC EFFICIENCY (INLET-TO-STATOR 0 LE) 0 III II !III -0.1 0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 11 (C) LE NORMAUZED AXIAL DISTANCE TE Figure 3: Comparison of Stage 0 blade exit radial profiles of total pressure, total temperature, and adiabatic Figure 6: Axial distribution of blade surface isentropic Mach efficiency at the aerodynamic design point. number near the hub for 1M2500+ Rotor 0.

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FIRST STAGE BLADE DESIGN CFO-80C2 and LM2500 first stage compressor blades are designs that use midspan shrouds. The LM2500+ first stage blade moved away from the shrouded design to a wide chord rotor design. Some of the key geometric and aerodynamic design para- meters for the first stage blade are presented in Table V. 1 1.2 • 1.1 _ Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 co 1.0 _ TIP 2 o 0.9 - z = 0.8 - 2 0.7 _ • 06 0. 0 cc 0.5 z 0.4 co - 0.3 _ Figure 7: LM2500+ ROW' 0 [AMC 02 1/1111111 0,4 -0.2 C 0.2 0,4 0,6 0.8 1.0 1.2 ZERO STAGE VANE DESIGN NORMALIZED AXIAL DISTANCE Stator 0 on the LM2500+ replaces the inlet guide vane on the (a) base machine but has aerodynamic characteristics similar to the 1.2 first stage vane in the baseline compressor. As mentioned before, 1.1 - the stator hub loading levels and inlet absolute Mach numbers • co 1.0 _ PITCH were controlled by using -15 degrees counter-swirl into the zero 2 stage blade. Table IV summarizes the geometric and aerodynam- o 0.9 _ ic design parameters for the vane. 0,8 _ 2 0.7 - Table IV: Zero Stage Vane Key Aerodynamic • Os 0 Design Parame ters cc 0.5 - 1•■ Parameters Zero Stage Vane Z 0 Number of Vanes 40 - 0.3 _ Tip, Pitch, Hub Chord (in.) 2.49, 2.01, 1.7 0.2 lip, Pitch, Hub Tmax/C .105, .081, .039 -0.4 -0.2 0.2 0.4 0.6 0.8 1.0 12 NORMALIZED AXIAL DISTANCE Inlet Absolute Hub Mach Number 0.95 Hub Diffusion Factor 0.55 (b)

In the paper on three-dimensional relief, Wadia and Reacher (1990) have shown that the need to align the inlet metal angle pre- HUB cisely with the skewed inlet flow angles at the endwall (at the risk of placing a significant spanwise twist gradient in the blade • 0.8 0 surface near the endwall) appears to be less than might be implied 2 0.7 from two-dimensional cascade analyses. While some recognition a.g 0.6' of the high air angles that exist at the endwalls was considered in 0 • 0.5 the design of the zero stage vane, extreme leading edge angle gra- dients were not required. Figure 8 shows the calculated three- 6 0.4 dimensional vane surface isentropic Mach number distribution at - 0.3 , , , , 111, the tip, pitch and hub. A shock-free diffusion at the hub was 0.2 1 -0.4 -02 02 0.4 0.6 0.8 1.0 12 accomplished and the ability of the stator to deliver the required LE TE swirl to the following rotor was verified by the three-dimensional NORMALIZED AXIAL DISTANCE analysis. (c) Figure Et: Axial distribution of Stage 0 vane surface isentropic Mach number.

6 Table V: First Stage Blade Key Aerodynamic Similar to the zero stage blade, this airfoil was also ruggedi- Design Parameters zed to improve FOD tolerance. Some amount of hub scalloping Parameters Stage 1 Blade was also done to accommodate the large thickness increase in the hub due to the elimination of the Midspan shroud and thicker tip Number of Blades 26 sections. Three-dimensional analysis was used extensively in the Tip, Pitch, Hub Solidity .95, 1.29, 1.86 design, and Figure 9 illustrates the isentropic Mach number dis- Aspect Ratio 2.337 tribution on the blade surface at the tip, pitch and hub immer- Inlet Relative Tip Mach Number 0.90 Pitch inlet Swirl Angle (degrees) 20 sions. As illustrated by the axial distribution of the blade surface Mach number at the hub and the midspan in Figure 9, most of the Chord (in.) 3.35 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 airfoil sections along the span were front loaded. However, the Tip, Pitch, Hub Tmax/C .035_090_13 sections locally near the tip were more aft loaded to reduce tip leakage and improve performance as has been demonstrated in 1.2 low-speed testing by Wisler (1985). As in the zero stage blade IMM 0.09 co 1.1 analysis, the three-dimensional analysis was used to verify that 2 the desired radial profiles of total pressure and efficiency were z achieved by the first stage blade. <0 0.9 - Figure 10 shows a photographic comparison of the current production first stage midspan shrouded blade and the new wide 0.8 chord LM2500+ first stage blade. The LM2500+ blade has about 0.7 35 percent lower aspect ratio and 60 percent more midspan chord but approximately the same midspan solidity relative to the first UJ 0.8 - co stage blade in the base machine. 05 111111111 -0.1 0 0.1 0.2 0.3 0.4 0.5 0.8 01 0.8 OS 1.0 ii LE TE NORMALIZED AXIAL DISTANCE (a)

IMM = 0.5

04 -0.1 0 0.1 0.2 0.3 0.4 0.5 0.8 0.7 0.8 0.9 1.0 11 TE LE NORMALIZED AXIAL DISTANCE Figure 10: Comparison between the LM2500+ wide chord (b) Rotor 1 and the LM2500 Rotor 1 with the midspan shroud. 1 aiiimorIMM = 0.92 FIRST STAGE VANE DESIGN The first stage vane design was only a minor change from 0.8 that in the CF6-80C2 core compressor. The design incidences on the LM2500+ Stator I were patterned after those from the CF6- U 0.7 80C2 Stator 2 as this would be the second stage of the LM2500+ 0 OS compressor. The vane surface isentropic Mach number distribu- tions on Stator 1 were similar to those reported for the zero stage 0.1 0.5 vane. The required inlet conditions into the next rotor stage, 0.4 which is identical to the CF6-80C2 airfoil, were verified by the -0.1 0 0.1 0.2 0.3 0.4 0.5 OS 0.7 0.8 0.9 1.0 11 LE TE three-dimensional analysis. NORMALIZED AXIAL DISTANCE (c)

Figure 9: Axial distribution of blade surface isentropic Mach number for Rotor 1.

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LM2500+ VARIABLE STATOR SYSTEM BO To ensure stall-free operation at part speed, the inlet guide if 70 vanes and the first six stators are variable in the LM2500+ engine. Mi cr 60 0 The method of scheduling the variable stators was changed in the tu LM2500+. Whereas the LM2500 uses a pair of lever arms O SO powered by twin actuators, the LM2500+ uses two torque shaft in 00 assemblies located 180 degrees apart to actuate the compressor g g 30 variable stator vanes. Each torque shaft, which is cylindrically shaped, is mounted to the compressor casing along the axis of the ei 20 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 engine with a bearing at each end to allow for rotation within a rn 10 forward and aft mounting bracket. A hydraulic actuator attached 0/ 0 to a lug near the front end of each shaft rotates the shaft. Additional lugs, one for each variable stator stage, are bolted -1 0 along the length of the shaft. As the shaft turns, a tumbuclde 5 10 15 20 25 30 35 40 45 50 bolted to each lug is pushed or pulled to rotate a multipiece, 360- STATOR 2 (MASTER) ANGLE (DEGREES) degree ring to which each stator vane lever arm is connected, Figure 12: LM2500+ compressor pretest variable stator opening or closing the VSV. The key benefit derived by using the gang relation. torque tube is its added flexibility to schedule each stage with stator angle schedules that are non-linear to each other. The variable stator schedule is controlled as a function of the TEST COMPRESSOR INSTRUMENTATION SUMMARY compressor inlet corrected speed, with the second stage stator The first engine with the LM2500+ compressor was tested as angle used as the reference. Stage characteristic data sets acquired a gas generator with a conic nozzle at GE Aircraft Engines test from core engine tests of the CF6-80C2 and the LM2500 com- facilities in Evendale with a Dry Low Emissions (DLE) combus- pressor were used to assemble a pitch line off-design model of the tion system. LM2500+ compressor with the new inlet guide vane and zero The overall compressor performance was measured using stage, and the mixture of airfoils from the CF6-80C2 and four five-element combined inlet total pressure and total temper- LM2500. This model was used to develop a pretest stator sched- ature rakes and three five-element rakes at the compressor dis- ule for the LM2500+. charge. Other performance measurements included the compres- The change in Stator 2 angle (pretest) from the design sor speed and inlet flow, which was measured with a calibrated nominal as a function of compressor corrected speed is shown in bellmouth with inlet rakes. All compressor bleed flows were mea- Figure 11. The change in the setting angles from design for the sured using calibrated pipes. inlet guide vanes and Stators 0, 1, and 3-6 are shown in Figure 12 lnterstage compressor instrumentation was extensive. Each as a function of the Stator 2 "master" angle. While the linear rela- stage in the compressor had leading edge total pressure and total tionship of Stators 0-6 with Stator 2 was retained from the temperature instrumentation. All the compressor airfoils were CF6-80C2 and the LM2500 experience, the inlet guide vane rela- instrumented with either flame-sprayed or thin film strain gages, tionship was highly non-linear with respect to the "master" and light probe instrumentation was included over Rotor 0 and stator angle. Rotor I. Two vane stem-mounted potentiometers per stage were used LAR2500+ COMPRESSOR STATOR SCHEDULE on all variable stators to monitor stator position during the test. 454s Trimmer motors were provided for all variable stators to individ- ually vary the stator settings for performance/aeromechanical 40 to optimization. tu tu 35 Tip clearance measurements were obtained over the rotors in cc stages 0, 1, 2, 6, 11, and 16. All the compressor airfoils in the iu 30 LM2500+ that were common to the base engine were from a lease w 25 pool engine (which had about 700 hours run time) including the Z 20 rear compressor case. The measured average clearance to blade 4 height, at design speed, varied between 0.7 to 1.%, except for ix Is 0 stage 16 which had significantly larger clearances (-5%). The CE 10 clearances on the new airfoils and on the airfoils from the lease pool engine were somewhat larger than one would expect in a • 5 new engine. I 2000 4000 6000 8000 10000 12000 COMPRESSOR TEST PERFORMANCE RESULTS INLET CORRECTED SPEED (RPM) Figure 13a shows the comparison of the measured and pre- :Figure 11: LM2500+ compressor stator schedule. dicted compressor operating lines with the pretest variable stator schedule. The distinct steps in the compressor operating line rep-

8 40 resent the different stages of operation of the DLE combustion

35 system, which uses compressor discharge bleed to control the combustion flame temperature. The compressor achieved the 30 design flow at the design speed, and the agreement between the 25 calculated and measured flow at speed with the pretest variable stator schedule as shown in Figure 13b was very good. The com- 20 pressor achieved its peak polytropic efficiency, slightly over 91 Er 15 percent, near its aerodynamic design point as shown in Figure 13c. The test data shown in Figure 13c is as measured, while the aj 10 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 cc pretest prediction line includes an estimated 0.5 points derate in 5 target efficiency due to compressor instrumentation, hardware quality and clearances. The individual performance for the blisk and that of the downstream compressor with the mixture of CF6-80C2 and LM2500 airfoils was separated in the following manner. The zero -10 0 20 40 60 80 100 120 140 160 180 200 220 stage performance was determined by using data from the inlet INLET CORRECTED FLOW (MS/SEC) rakes and the zero stage, vane-mounted probes. In spite of the (a) large clearances and the significant amount of ruggedization in the 220 transonic Rotor 0, the IGV-Rotor 0 combination achieved a peak polytropic efficiency of 90.5 percent, which was 2.8 percent better 200 E.; than design intent. The downstream compressor performance with 180 the mixed airfoils, as determined by using the zero stage, vane- mounted probes and the exit rakes (and referred to as the "rear _1 160 block" in this paper), also achieved the desired match point for the 0 140 flow and efficiency goal relative to the baseline compressor. The measured radial profile of total pressure and temperature at Rotor 0 exit was in good agreement with the three-dimensional calculations and the measured efficiency was slightly higher than Predicted Test Data the calculations near the tip. Aeromechanically, the first two new stages performed per design intent over the entire operating range including vane off- — 40 . schedule operation. However, among all the other compressor blades, strain gage signals on Rotors 2, 3, and 4 indicated higher 20 than desired stresses between 83 to 90 percent design speed. A 5000 6000 7000 8000 9000 10000 INLET CORRECTED SPEED (RPM) small variable stator schedule adjustment, using the trimmer (D) motors on the variable stators, provided an acceptable solution to .92 the aeromechanical stress issue. The loading on these rotors was reduced by a combination of the change in their upstream and .90 downstream stators, resulting in a significant reduction in aero- .88 mechanical response. As a result, the IGV through Stator 2 were closed 3 degrees from nominal (only at part speed), and Stators 5 z .86 LU and 6 were opened by 3 and 5 degrees from nominal, respective- .84 ly. The loading relief on Rotors 2, 3, and 4 came as a consequence U. LU of pushing the load forward onto the zero and first stages and onto .82 o the downstream side loading up the middle stages (Rotors 5, 6, a. 0 cc •Bo and 7) of the compressor. The schedule adjustment did not

-I .78 adversely affect compressor operability as supported by data later 0 • in this paper. .76 The strain gages on all the compressor vanes indicated low stresses. However, the high aspect ratio inlet guide vanes, which .74 have the largest travel with compressor speed, exceeded accept- 72 able design stresses between 93 to 101 percent design speed. The 0 20 40 so 80 100 120 140 160 180 200 220 INLET CORRECTED FLOW (LESS/SEC) 1GV stress issue was resolved by a variable stator gang closure of (C) 3 degrees at high speed. This final production ganging of the stators to the torque tube Figure 13: LM2500+ compressor performance with pretest was defined and mechanically verified while on test. This experi- variable stator schedule. mentally determined master stator schedule as a function of inlet

9 •

corrected speed, labeled as the "stress rig" in succeeding para- efficiency dropping to 89.5 percent due to the change in the vari- graphs, is shown in Figure 11. able stator schedule. 1.50 COMPRESSOR PERFORMANCE WITH "STRESS RIG" STATOR SCHEDULE O "5 0 The optimized stator schedule to reduce the stresses in the ▪ 1AD inlet guide vanes and the front rotors reduced the inlet corrected La 1.35 0 z 0 flow by 6 percent at design speed. The loss in flow was retrieved u, 1.30 by increasing the compressor speed by 1.4 percent with about 0.3 Lb sr 1.25 0 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 percent loss in peak overall compressor polytropic efficiency. o. 0 1.20 Figure I 4a shows the reduction in flow at speed along the com- cc 0 pressor nominal operating line with the "stress rig" variable stator g 1.15 - 0 0 schedule. The compressor polytropic efficiency along the engine CC 1.10 operating line versus compressor inlet corrected flow is shown in O 1.05 Figure 14b. The compressor achieved a peak polytropic efficiency of 100 90.8% at 73.6 kg/s (162 lbs/sec) at 19.8:1 pressure ratio. 70 80 90 100 110 120 130 140 150 160 170 180 190 INLET CORRECTED FLOW (LBS/SEC) 180 (a) Ic3 160 .95 a m 140 .9a

CIENCY 0 0 0 3 120 0 .85 - / 11. 100 EFFI 0 0 Ill 60 .80

60 ▪ .75 0 CC 0 ADIABATIC .70 0 2 .65 ROTOR 0 6000 7000 8000 9000 10000 IGV- ...... INLET CORRECTED SPEED (RPM) 60 70 80 90 100 110 120 130 140 150 160 170 180 190 (a) INLET CORRECTED FLOW (LBS/SEC) .94 (b) Figure 15: LM2500+ IGV-Rotor 0 performance. .89 The downstream compressor performance with the mixed IENCY airfoils was determined by using the zero stage, vane-mounted FIC .84

EF probes and the compressor exit rakes. Figure 16a shows the com-

IC parison between the base LM2500 compressor operating line and .79 the measured distribution of pressure ratio with flow for the "rear block" on the LM2500+. Note that the corrected flow for the

OLYTFIOP .74 LM2500+ "rear block" was calculated using the zero stage, vane. P mounted, leading edge measurements for total pressure and total temperature while that for the base LM2500 compressor truly rep- 69 o 20 40 60 80 100 120 140 160 180 200 resents the compressor inlet conditions. The LM2500 (small INLET CORRECTED FLOW (LBS/SEC) symbols) and the LM2500+ (large symbols) were both run with (b) the DLE combustion system and hence the "stepped" (for com- bustor staging) compressor operating line. Data for the base Figure 14: LAA2500+ compressor performance with "stress rig" variable stator schedule. LM2500 engine with the Single Annular Combustion (SAC) system is also shown as a solid line without symbols in the figure. The zero stage performance with the "stress rig" was deter- Figure 16b shows the corresponding comparison of efficiency at mined using data from the inlet rakes and the zero stage, vane. flow for the base LM2500 compressor and the LM2500+ com- mounted probes. Figure 15a shows the pressure ratio at flow pressor "rear block." The LM2500+ "rear block" shows approxi- achieved by the IGV-Rotor 0 combination along the engine oper- mately 0.8 points improvement in peak efficiency, which is in ating line. The corresponding efficiency at flow for the IGV-Rotor agreement with the estimated improvement in efficiency with the 0 combination is shown in Figure 15b, with the peak polytropic custom-tailored CF6-80C2 airfoils. While not a technical require-

10 ment, these airfoils also provided a significant improvement in 45 performance at lower corrected flows. 40 22 35

30 cc zs 0 0 20

0 15 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 10

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0 3000 4000 5000 6000 7000 8000 9000 10000 INLET CORRECTED SPEED (RPM) Figure 17: 1..M2500+ compressor open stator stall test 30 40 50 60 70 80 90 100 110 120 130 140 150 160 results. REAR BLOCK CORRECTED FLOW (LBS/SEC) CONCLUDING REMARKS (a) .935 The aerodynamic design of the 17-stage, 23.3:1 pressure ratio single spool axial compressor for the LM2500+ gas turbine tj.eas has been reported. The new zero stage and stage 1 compressor E. blades and vanes were designed using three-dimensional compu- 0 .835 tational methods. Advanced compressor airfoils from the CF6-80C2 compressor were also integrated successfully with the 0 baseline LM2.500 compressor airfoils. The compressor achieved E .785 a peak polytropic efficiency slightly over 91 percent and met all 0 G. its design operability objectives. NC .735 In spite of the significant ruggedization of the front-stage 0 CO airfoils operating in the transonic flow regime, advanced design I. tools were able to provide accurate guidance during the design cc phase to achieve the level of compressor efficiency required for 635 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160 the successful field implementation of the LM2500+ at its full REAR BLOCK CORRECTED FLOW (LBS/SEC) mature power rating. Production engine data have duplicated the (b) performance predictions based on the development engine. Figure 16: LM2500+ "rear block" performance. ACKNOWLEDGEMENT The compressor was not throttled to stall at these high pres- The authors would like to thank GE Aircraft Engines for sure ratios. Stall-free compressor performance was demonstrated permission to publish this paper. along a 7 percent high operating line, however, by adding flow blockers (-tabs") to the exhaust nozzle to reduce the area. REFERENCES Open stator stall testing to determine compressor "open Adkins, GO., and Smith, L.H., 1982, "Spanwise Mixing in Axial Flow Thrbomachines," stator stall margin" (defined as the difference between the sched- ASME Journal of Engineering for Power, Vol. 104, pp. 97-110. uled VSV angle and the VSV angle at stall) was conducted on the first development test engine. Holding constant speed, the vari- Dunavant, IC., "Cascade Investigation of a Related Series of able stator gang was opened in one-degree intervals until the com- 6 Percent Thick Guide Vane Profile and Design Charts," pressor stalled. Figure 17 shows the open stator stall test results. NACA-TN-3959. The VSV gang based on Stator 2 is shown plotted versus the com- pressor inlet corrected speed. The stator position at which the Eisenberg, B., 1993, "Development of a New Front Stage for an compressor stalled as the VSV gang was opened is also shown in Industrial Axial Flow Compressor," ASME Paper No. 93-GT-327. Figure 17. At 90% design speed and beyond, the VSV gang was Farmer, R., 1994, "GE Launches LM2500+ Rated at 39 MW and opened to demonstrate the required margin without stalling the 38% Thermal Efficiency," Gas Turbine World, May/June 1994. compressor. The LM2500+ compressor demonstrated over ten pp. 24-32. degrees margin at 9,000 rpm and over five degrees at 5,000 rpm, which is adequate. The actual margin at the open stop was not Janssen, M., Zimmermann, H., Kopper, F., and Richardson, J., determined but is enough for stall-free operation of the LM2500+ 1995, "Application of Aero-Engine Technology to Heavy Duty compressor. Gas Turbines," ASME Paper No. 95-GT-133.

11 Jennions, I.K., and Turner, M.G., 1993, 'Three-Dimensional Sehra, A., Bettner, J., and Cohn, A., 1991, "Design of a High Navier Stokes Computations of Transonic Fan Flow Using an Performance Axial Compressor for Utility Gas Turbine," ASME Explicit Flow Solver and Implicit k-e Solver," ASME Journal of Paper No. 91-GT-I45. Turbomachinety, Vol. 115, pp. 261-272. Smed, J., Pisz, F., Kain, J., Yamaguchi, N., Umemum, S., 1991, • Kashiwabara, Y., Katoh, Y., Ishii, H., Hattori, T., Matsura, Y., and "501F Compressor Development Program," ASME Paper No. Sasada, T., 1990, "Developments Leading to an Axial Flow 91-GT-226. Compressor for a 25MW Class High Efficiency Gas Turbine," ASME Paper No. 90-GT-238. Stringham, G., Cassem, T., Prince, T., and Yeung, P., 1998,

"Design and Development of a Nine Stage Axial Flow Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1999/78583/V001T03A042/4215696/v001t03a042-99-gt-210.pdf by guest on 02 October 2021 Katoh, Y., Kashiwabara, Y., Ishii, H., Tsuda, Y., and Yanagida, M., Compressor for Industrial Gas Turbines," ASME Paper No. 1993, "Development of a Transonic Front Stage of an Axial Flow 98-GT-I40. Compressor for Industrial Gas Turbines," ASME Paper No. 93-GT-304. Valenti, M., 1998, "Luxury Liners Go Green," ASME Mechanical Engineering, July 1998, pp. 72-73. Klapproth, J.F., Miller, ML., and Parker, D.E., 1979, "Aerodynamic Development and Performance of the Wadia, A.R., and Beacher, BY, 1990, "Three-Dimensional CF6-6/LM2500 Compressor," AIAA Paper No. 79-7030. Relief in Blading," ASME Journal of Turbomachinery, Vol. 12, No. 4, pp. 587-598. Koch, C.C., 1981, "Stalling Pressure Rise Capacity of Axial Flow Compressor Stages," ASME Journal of Engineering for Power, Wadia, A.R., and Law, C.H., 1993, "Low Aspect Ratio Transonic October 1981. Rotors: Part 2 — Influence of Location of Maximum Thickness on Transonic Compressor Performance," ASME Journal of • Koch, CC., and Smith, L.H., 1975, "Loss Sources and Turbomachinery, Vol. 115, pp. 226-239. Magnitudes in Axial Flow Compressors," GE Aircraft Engines Technical Report No. R75AEG344. Wadia, A.R., and Copenhaver, W.W., 1996, "An Investigation of the Effect of Cascade Area Ratios on Transonic Compressor Leuven, V.V., 1994, " Incorporated Taurus 60 Gas Performance," ASME Journal of Turbomachinery, Vol. 118, Turbine Development," ASME Paper No. 94-GT-I15. October 1996, pp. 760-770.

Rocha, G., Saadatmand, M., and Bolander, G., 1995, Wisler, D.C., 1985, "Loss Reduction in Axial-Flow Compressors "Development of the Taurus 70 Industrial Gas Turbine," Through Low Speed Model Testing," ASME Journal of ASME Paper No. 95-GT-41I. Engineering for Gas Turbines and Power, Vol. 107, pp.354-363. Scalzo, A., and Mod, Y., 1988, "A New 150 MW High Efficiency Heavy Duty Combustion Turbine," ASME Paper No. 88-GT-162.

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