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A Cost-Effective Performance Development of the PT6A-65

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A Cost-Effective Performance Development of the PT6A-65 E AMEICA SOCIEY O MECAICA EGIEES 864 4 E. 4 St., Yr, .Y. 400 e Sociey sa o e esosie o Saemes o oiios aace i aes o i is- cussio a meeigs o e Sociey o o is iisios o Secios o ie i is uicaios iscussio is ie oy i e ae is uise i a ASME oua aes ae aaiae om ASME o iee mos ae e meeig ie i USA Copyright © 1985 by ASME Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/IGT1985/79429/V001T02A015/2477464/v001t02a015-85-igt-41.pdf by guest on 27 September 2021 A CtEfftv rfrn vlpnt f th 6A6 rbprp Cprr SUKASA YOSIAKA nd KEE S. UE Aeoyamics eame a Wiey Caaa Ic ABSTRACT B = T0/288.3 (°K) : enthalpy rise coefficient = ip/ri The largest member of the PT6 turboprop engine family, the PT6A-65, was : flow coefficient = Cm/U developed in the early 1980's and went into production in September 1982. The : pressure rise coefficient = (JOH 0 )/U 2 compressor for this engine consisted of four new axial stages combined with an existing centrifugal stage on a single shaft. This paper gives a brief description of Subscripts the studies leading up to the choice of the compressor configuration and a more c : choking detailed examination of the development of the chosen compressor to the is : isentropic required performance level. m : meridional component The development of this compressor presented a two-fold technical max : maximum challenge. Firstly, the limited space in the small compressor gas path did not p : peak efficiency point permit the effective use of conventional total pressure and temperature probes : surging point for performance evaluation. Secondly, the short time available for development o : engine inlet excluded some attractive corrective measures such as the redesign of some of the 3 : diffuser exit axial blade rows because the time required would have jeopardized the meeting of the tight development deadline. The first problem was overcome by a combination of limited wall static pressure measurements and an extensive use INTRODUCTION of numberical flow analysis codes. This approach proved to be quite cost- effective. The second was solved by the adaptation of an existing fully The first PT6 turboprop engine, PT6A-6, was marketed in 1962. By the analytically-designed research axial stage to the first stage position in the axial mid 1970's this ancestor had sired an extensive series of turboprop engines compressor. covering a take-off thermal power range from 450 shp to 1120 shp (gearbox limited) under sea level standard day conditions. In 1974, Pratt & Whitney NOMENCLATURE Canada Inc. (P&WC) was forecasting a demand in the 1980-1990 time period for a turboprop engine in the 1200-1300 shp range for high-speed, high-altitude A cross-sectional area at the leading edge of the 1st stage rotor (m2) executive aircraft applications. A series of turbomachinery variations were C absolute velocity (m/sec), or chord length (mm) studied in an effort to satisfy the predicted performance requirements. In the D F diffusion factor case of the compressor for this engine the number of possible configurations had i absolute total enthalpy change across the compressor or turbine been reduced to three by the spring of 1978. These were (A) the addition of a (Joule/ kg) zeroth stage to an existing compressor; (B) the combination of a newly designed MCA multiple circular arc four stage axial compressor with an existing centrifugal stage (4A + IC); (C) a m mass flow (kg/sec) completely new 3A + IC compressor. Configuration A was eliminated in spite A = (ria-fri) (kg/sec) at a constant speed of its low technical risk and low development cost advantages due to insufficient absolute total pressure (kg/ern 2) performance improvement over the current PT6 models. Configurations B and Ps static pressure (kg/cm-) C were predicted to be about equal with respect to performance. However, the PR pressure ratio former was considered to be a lower risk design than the latter due to the lower R rotor, or radius (mm) pressure ratio per stage for the axial stages and the use of the existing high S stator performance centrifugal stage. In addition the 4A + IC configuration (B) was S/M surge margin = HPR/MV-0/6)] s/[PR/M/0/6)], — I x 100 ,70 expected to have a higher uprating potential than the 3A + IC (C). T absolute total temperature (°K) Consequently, the 4A + IC configuration was the final choice . blade thickness (mm) This paper begins with a brief description of the aerodynamic design of the U rotational speed at mean radius (m/sec) four axial stages for this new compressor. This is followed by a more detailed X axial length (mm) discussion of the performance development of the compressor. The blade or vane stagger angle (deg.) development was technically challenging for the following reasons: P0/1.033 (kg/cm 2) I Conventional total pressure and temperature probes for the evaluation of compressor adiabatic efficiency compressor stage performance could not be used effectively in this unit Presented at the 195 eiig Ieaioa Gas uie Symosium a Eosiio eiig eoes euic o Cia — Seeme 1-7 195 because of the small gas-path dimensions (see fig. 1). They had the = 1 undesirable effects of limiting flow at high speed and of introducing excessive losses. 7 1 2. Due to the tight development schedule, it was impossible to undertake a redesign of the compressor without imperiling the meeting of the „- 9 development deadline. Even the time allowed for the optimization of stage matching was severely limited for the same reason. E 9 -631' Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/IGT1985/79429/V001T02A015/2477464/v001t02a015-85-igt-41.pdf by guest on 27 September 2021 E 1.6 o oo eaie Mac ume "a is" E 1.4 Sao asoue Mac ume "a u" 6 [150- 1.2 1.0 eicio 0.8 R -es esus 0 6 1 Ei ° 4 - 1 oi Imee 9 97 J (WA/(/A esig R4 S4 S1 9 = 1 (% R1 45 39 2 50 3 aes •7 15 blades 0 99 mm) 0 50 100 15 200 93 1 770 97 . 4 (inches)0 2 4 X 6 8 cc 97 19 99 igue 1 Meiioa aia comesso gas a (oigia esig 60 93 c_ 97 As a result of these constraints the compressor performance development 87.2 was limited to gas generator tests with minimum gas-path instrumentation. 5 AXIAL COMPRESSOR DESIGN 75 80 5 9 95 1 15 (iogo/o / (.oA/o esig (° The four axial stages of the new compressor were designed during the second half of 1978 using the well established P&WC axial compressor design igue Comesso ma (coig 1 system. The design point pressure ratio for the axial stages was 4.26:1. The compressor had a "front loading" type of work distribution which corresponds approximately to a constant enthalpy rise per stage. A maximum design The philosophy of no variable vanes, which is one of the traditional diffusion factor D, of 0.45 was selected to ensure an adequate surge margin features of PT6 compressors ensuring high engine reliability, was maintained for without variable geometry. This limit combined with the relatively high stage this new design. However, an interstage bleed arrangement located between the loading implied in the design pressure ratio resulted in a high tip speed design. fourth stator and the impeller, as on the other PT6 models, was used and a jet- Consequently, the inlet relative flows to the first three rotors are transonic with flap system was added to the intake case. These features have been more than the tip relative Mach number being 1.49, 1.17 and 1.05 for the first, second and adequate to avoid possible mismatch problems between the front and rear stages third stage rotors, respectively. The meridional gas path geometry of the axial at start-up and idle. They are both shut off at all power rating points. stages, (fig. 1), displays an essentially contant average radius. The correponding design point mean line Cm distribution was also nearly contant. Table 1 COMPRESSOR PERFORMANCE DEVELOPMENT contains the design values of various aerodyamic and geometric parameters for the axial stages. By the time the engine development was embarked upon in late 1979 two customers had decided to use this engine, designated PT6A-65, on their new commuter aircraft. In addition, one of them had a plan to develop an executive ae Rotor Stator Rotor Stator Rotor Stator Rotor Stator version of the new airplane. This implied that the compressor, originally ow 1 1 2 2 3 3 4 4 designed for high-speed, high-altitude applications, was now required to provide Ie ai age eg 15 353 60.8 1 59 35 5 7 high performance down to 9 61 of design speed, N 1 , for hot day take-off, while Ei ai age eg 5 11 59 1 9 13 42.0 5 maintaining the low sfc capability up to at least 105% N1, 1 for high altitude Ie mea age eg 5 15 577 77 5 7 571 9 cruise. This had to be accomplished without variable geometry. Ei mea age eg 55 35 5 2.8 11 39 3 1 Soiiy 19 11 19 1 19 1 15 13 Test of Baseline Compressor (Config. 1) ickess/co 51 .064 7 1 5 1 59 .060 aco 39 77 39 359 .422 3 .426 33 In November 1979 the first compressor map was taken between 89.7% and A ages om aia 100% N 1 on a gas generator which had been allocated to the compressor development program.
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