The Effect of Reaction on Axial Flow Compressor Performance

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THE AMERICAN SOCIETY OF MECHANICAL ENGINEERS 345 E. 47th St., New York, N.Y. 10017 The Society shall not be responsible for statements or opinions advanced in papers or discussion at meetings of the Society or of its Divisions or Sections, 94-GT-456 or printed in its publications. Discussion Is printed only If the paper is published in an ASME Journal. Papers ere available from ASME for 15 months altar the meeting. Printed in U.S.A. Copyright © 1994 by ASME

THE EFFECT OF REACTION ON AXIAL FLOW COMPRESSOR PERFORMANCE Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021

C. D. Farmakalides OPRA Hengelo, The Netherlands

A. B. McKenzie and R. L. Elder School of Mechanical Engineering Cranfield University 111111111111111111111111 Bedford, United Kingdom

ABSTRACT GREEK SYMBOLS This paper describes a study of the effects of design-point a Flow angle reaction on axial flow compressor performance. Particular am Mean flow angle attention is given to differences in stable operating range and Blade angle overall efficiency. Design of an 80% reaction, zero pre-whirl, 8 Blade deviation angle blading is presented together with a discussion on the applicability 0 Blade camber of currently available design correlations, mostly derived from 50% Blade stagger angle reaction binding tests, to such high reaction blading. Experimental a Blade solidity - 11(5/C) data obtained from tests carried out on a low speed 3-stage axial (1) Flow coefficient - C/U flow research facility, at Cranfteld Institute of Technology (now Cranfield University), using an existing 50% reaction blading and INTRODUCTION the new 80% reaction blading indicate that high reaction designs A stage velocity diagram may be specified by the choice of three can result in improved operating range at no loss of efficiency. parameters. Modem axial flow compressor blading differs only Tests carried out include performance measurements for each of slightly with respect to two of these parameters, namely work and the two bladings at various stator stagger settings and include inter- flow coefficients, so the diagram can be characterised by the third, row radial traversing at flow conditions near optimum and stall. the degree of stage reaction. For the last fifty years, the effect of reaction on compressor performance has been somewhat NOMENCLATURE controversial and in the early days in Britain the designs tended CO AVDR Axial velocity density ratio concentrate around 50% while German designs tended to favour Blade chord high reaction around 90% as used in the Jumo 004 engine. C, Axial component of absolute velocity Arguments may be found by Howell 11945]. Seippel 119401. which D.F Diffusion factor consider Mach number limitations to support 50% reaction. There Stagnation (total) enthalpy are arguments, however, Cumpsty [1989] and Casey [19871, Blade incidence angle suggesting that losses are mainly insensitive to reaction and Hans- Compressor inner diameter, see Davis [1973] Otto andHeinrich [1984] claims a larger stable operating range for Compressor outer diameter high reaction blading. Blade pitch (space) Industrial compressors an required to have the greatest possible Space-chord ratio range of mass flow at constant speed and pressure ratio. This Blade thickness range is usually achieved by using variable stagger stators. Tip clearance Variable geometry compressors, however, are considerably more Blade thickness to chord ratio expensive than fixed geometry compressors and there are clear Blade speed benefits to be achieved by minimising variable components and

Presented at the International Gas Turbine and Aeroengine Congress and Exposition The Hague, Netherlands — June 13-16,1994 introducing high reaction blading if the penalties are small. low Mach number (<0.7) applications and is compatible with the The purpose of this study has been to evaluate the effects of existing design. reaction on compressor performance with particular attention to The work coefficient, niii/U 2 = 0.4, is considered to be typical of

optimum efficiency and stable range of operation. Although this current trends in modem high designs (see for example the papers Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 area has attracted attention in the past, most of the work has been of Dong et al [1987] and Dring et al [1986] ). The flow unpublished and it is not clear whether properly designed high coefficient, CIO = 0.6, was selected from considerations of reaction blading can be competitive (on range and efficiency) with solidity, De Haller number, diffusion factor together with other the lower 50% reaction blading for which abundant design data is factors and is also believed to be representative of current designs. available in the open literature. The 80% reaction design was so chosen because this design has The work carried out during the study Farmakalides [1992], zero-whirl at stator exit for all radii (a well known form of design) involved the following areas which are discussed in the following and therefore has the practical advantage of requiring no I.G.V's sections. for a variable speed machine (and zero camber I.G.V's for variable (i) Design and manufacture of 80% reaction blading. stator constant speed applications) and the last stator requires no (ii) Experimental investigation of two kinds of blading with more deflection than the others. different stage reactions, each at various stator stagger settings. (iii) Evaluation of results BLADE LOADING CONSIDERATIONS The expression developed by Lieblein [1953] to evaluate the DESIGN OF BLADING diffusion factor (D.F.) is employed in its incompressible form. Reviews describing semi-empirical and computer aided Lieblein suggests values for D.F. not greater than 0.6 with typical approaches to axial flow compressor design have been published design values of 0.45 and the design met those recommendations. among others by Gostelow et al [1969], Howell [1945 and 1948], Horlock [1973], the NACA/NASA staff, Johnsen et al [1965], BLADE PROFILE DESIGN Serovy [1966] and AGARD [1981 and 1989]. These methods Reviews which describe semi-empirical methods of design have were generally adopted for the design of the 80% reaction blading been published by Carter [1950], Lieblein [1960], Horlock [1973], for this study (the 50% reaction blading existing previously). Serovy [1966] and many others. Tests carried out on single and The designs had the following parameters at a reference diameter multi-stage axial flow compressors have introduced empirical of 335.3 mm: factors which relate cascade performance to that of an actual compressor. In the case of McKenzie 11980 and 1988], a design Reaction 80% 50% method has been developed directly from compressor test data Work coefficient AN/U2 0.4 0326 Two methods of design were adopted and are referred to in this Flow coefficient Ci/U 0.6 0565 project as the McKenzie and NASA methods. The so called Whirl velocity at rotor inlet 0 30° NASA method does not follow exactly the recommendations of SP-36 [Johnsen et al 1965] but is in fact a combination of well 50% Reaction Bladinq established cascade correlations including Lieblein [1960], Carter The existing blading for this compressor is of free vortex design [1950] and Howell [1945]. with reaction varying from around 31% at the inner diameter The blade profile for the new designs is of C-4 thickness (ID.), to around 66% at the outer diameter (0.13.). The blades distribution on circular arc mean camber line with 10% maximum have a C-4 profile thickness distribution on circular arc camber thickness to chord ratio (t/C). To aid CAD-CAM manufacturing line with a maximum thickness to chord (t/C) ratio of 10%. Blade of the new blading the thickness distribution for C-4 is obtained chord is constant along the span at 3048 mm, to give an aspect from the relations given by Calvert [1988] and repeated in ratio (A.R.) of 2.0. The blades operate with a tip clearance (c) of Farmalcalides [1992]. 0.6 mm (0.15% of 0.D). Various aerodynamic and geometric design parameters are given S/C Ratio Selection in Table 1. Blade geometrical angles to meet the above conditions The starting point for blading design is the appropriate selection were calculated by the compressor designer using recommendations for space-chord ratio (S/C) which has been influenced by two of Roxbee-Cox [1946]. 'rules of thumb' as follows: (i) Over-blading is avoided by setting a minimum restriction 80% Reaction Bladinq on S/C ratio of 0.5. The new blading for the compressor is of free vortex design with (ii) Design point incidence is restricted to within ± 5°. In degrees of reaction varying from 71% at the inner diameter (I.D.) cases where incidence drops below -5°, the chord length to 86% at the outer diameter (OD.). The profile shape and tip is increased to reduce S/C and camber and move the clearance for this blading are as for the existing 50% reaction incidence to more positive values. build. The choice of free vortex blading is generally successful for

2 McKenzie Design Method Choice of S/C Ratio. Howell [1945] developed a blade loading McKenzie [1980] developed a design method based on tests of correlation in the form of curves of nominal fluid deflection LS vs a 4 stage low speed compressor. The relation for stagger (t) fluid outlet angle Cr., for various S/C ratios. The plot is used to suggested in his paper may be used to position the blade for indicate a SIC ratio selection compatible with an acceptable maximum efficiency. In order to increase stall margin, particularly loading to avoid boundary layer separation. for low stagger binding. McKenzie [1990] suggests the relation: Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021

Tan t Tana. - .15 (1) NASA design procedure. Selection for SIC is initially made using Howell's plot. The expressions derived for i„„ and S were combined to derive values for the camber angle, as follows: The relation for deviation developed by McKenzie [1980] is given by: 41- al -WA (K A Mo (7) 6 a (1.1 + .310) (8/0 10 (2) 1. —(m — n)

The relation for camber may therefore be developed by BLADE TO BLADE (B.T.B.) ANALYSIS combining equations (1) To assist the blade element selection, an analysis code was and (2) with circular arc mean line relations to give: implemented during the design period. The code is a 2-D finite element calculation applied to the full potential equation, Weber et Tan' (Tana. - .15) - a2 + 1.1 (S/C) 113 e (3) al [1987]. Viscous effects are considered in the code by using 3 - .31 (RP° McNally's [1970] integral boundary layer method coupled iteratively to the inviscid calculation by adding the displacement thickness to the metal blade contour. McKenzie [1988] considered a diffuser performance analogy to The code was employed primarily for analysing the effect of o relate efficiency to the ideal static pressure rise coefficient C . His incidence on the blade surface velocity distributions and also to correlation takes the form of a plot, where efficiency contours are provide some information for the fluid outlet angle obtained by o drawn for C versus S/C ratio values. The spline of these contours satisfying the Kam condition at the trailing edge. defines the relation for optimum S/C ratio given by:

S/C 9 (.567 - Cos) (4) FINAL DESIGN PROCEDURE The results from the analysis carried out using the NASA. McKenzie and blade to blade calculations were compared in order McKenzie's method may be used to analyse a given blade to provide an insight to the best way to proceed for final design. geometry by rearranging the equations to obtain the relation for Each of the two methods considered failed to account for the optimum incidence i ,, : complete range of camber and stagger angles encountered for both the rotor and stator blades. McKenzie's deviation rule is - Tan -1 1 2 Tai(, • I) + .15 - Tan(5 2 +6) - (5) questionable for low stagger,high camber sections while the NASA deviation rule is questionable for low camber, high stagger sections. Both Lieblein's reference incidence rule and McKenzie's where 5 is given by equation (2). stagger rule (which positions the blade to operate at optimum incidence) are found adequate for the blade sections considered. NASA Design Method Final design of the blades was achieved as follows: Incidence Correlation. The study report by Leiblein [1960] (i) Use Lieblein's correlation for both rotor and stator bladings to was used to define the reference incidence, i s. The analytical obtain an indication of the reference incidence i n,. expressions for Lieblein's correlation derived by Davis (1973) were (ii) The final choice for design incidence is obtained by employed. reducing the incidence from the reference incidence until a velocity Deviation Correlation. Reference is made to the relation for distribution is obtained (using B.T.B. calculations) where the deviation angle (8) proposed by Carter [1950] where: pressure side Mach number is never higher than the suction side (6) and never peaks in such a way as to imply a high diffusion rate from leading to trailing edge. A guide as to the desirable velocity distributions around blade sections was obtained from the In this study m is calculated using the analytical expression recommendations of Sauna [1988]. proposed by Davis [1973]. (iii) Use Carter's modified deviation correlation for stator designs and McKenzie's for rotor designs.

3 (iv) Use SIC ratios for rotor design as suggested by McKenzie assumption7 - was employed to relate the stage static pressure rise to and Howell's recommendations for stator blading design. the stage total pressure rise. Casey [1987] developed a mean-line prediction code (AXCHAR) Mass flow measurements were facilitated using the wall static for estimating the performance of axial flow compressors over the tappings at venturi inlet and throat planes - Ballal [1968] suggests full stable operating range from choke to stall. The method is a 1- an accuracy of ± 0.9% for the absolute mass flow measurements. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 D analysis of the velocity vectors at the root mean square (r.m.s.) Torque measurements were facilitated by a calibrated load cell radius with empirical relations for determining the losses and connected to the motor stator which was free to experience the operating range. Casey [1988] modified the code to extend it to reaction due to the motor's rotor. During operation the indicated predict a radial flow distribution within a repeating stage load varied by ± 0.3%. • assumption. The code may be applied in design or in analysis The shaft speed, measured to within 1 rpm, was set to within ± mode. 0.66% of the stated value during performance measurements and The final blading design was carried out using AXCHAR in the held to within ± 0.15% during traverse measurements. design mode. The resulting aerodynamic and geometric parameters Overall efficiency was limited due to the uncertainty in mass are indicated in Table 2. Because of a requirement for the test rig flow but repeatability was within ± 0.5 points of efficiency. to have the same number of blades per row as the existing 50% Traverse measurements for total pressure and flow angle were reaction blading, it was necessary to increase the true chord of both carried out using a calibrated 3-hole probe connected to a rotors and stators compared to the existing blading. Since the micromanometer. Radial traverse measurements were obtained for blade heights are the same, a reduction of blade aspect ratios was given stations of the probe positioned at stator blade mid pitch. unavoidable for the 80% reaction blading. Due to the increased All wall statics were measured using an inclined alcohol filled rotor blade stagger angles, its axial chord was only increased by manometer and were averaged around the circumference. The 7% and the reduction of its aspect ratio may be said to be a natural radial variation of static pressures were calculated on the advantage of a high reaction design. In the case of the stator this assumption of simple radial equilibrium. is not the case but its contribution to the stage static pressure rise is so low that the influence of its reduced aspect on stage stall is EXPERIMENTAL STUDY likely to be small. For each build, overall performance measurements are presented as mean stage characteristics of pressure rise and work coefficients TEST PROCEDURE and efficiency plotted against the flow coefficient. The re- The rig used for the experimental analysis is described in detail staggering of the stator blades changes the degree of reaction the by Farmakalides [1992]. The compressor consists of 3-stages of latter increasing with reducing stator stagger. 0.7 diameter ratio with provision for IGV's. The test program consisted of : Comparison B1, B2 and B3. The overall performance (i) Overall, stage and individual blade row performance characteristics of the three builds using the 50% reaction blading measurements is presented in figure 1. The perforrnance of the builds using the (ii) Interstage radial traversing at flow conditions near optimum 80% reaction blading is presented in figure 2. These figures and stall clearly indicate the shift in optimum flow coefficient as the stator (iii) Measurements repeated for stator stagger settings of design, stagger is changed as would be expected. design+10' and design-10°. Six builds have been tested and are identified as : 50% Reaction Bladinq tests Build 1 50% reaction blading; stators set at design stagger angle. A simple calculation based on mid-diameter will demonstrate the Build 2 50% reaction blading; stators set at design + 10° stagger change in optimum flow coefficient as stator stagger is changed. angle. The calculation is as follows:- Build 3 50% reaction blading; stators set at design - 10° stagger The flow coefficient (4)) is related to the flow angles by: angle. 1. Build 4 80% reaction blading; stators set at design stagger angle. 4) (8) Build 5 80% reaction blading; stators set at design + 10° stagger Tan a 1 + Tan a 4 angle. Build 6 80% reaction blading; stators set at design - 10° stagger From Build 1 for (I) = 0.530 and a, = 29° (obtained from angle. traverse measurements), a, is calculated using the equation above Measurements were obtained by running the compressor at 3000 as 53°. It is argued that the rotor blade optimum inlet flow rpm = 63.84 mis) conditions (a,) will be the same for the three builds. For the case a, will be some 10° higher (390) and for = Performance calculations were based on wall static pressures, of the second build, mass flow and torque measurements. The repeating stage 39° the flow coefficient at optimum is now calculated as 0.468.

4 This compares within 3% of the observed value. For build 3, a, General Discussion- all builds tested is now taken as 19° and the optimum flow coefficient is calculated The success of using variable geometry in efficiently producing as 0.598 which is within about 4% of that observed. The a constant pressure rise at a range of mass flows is indicated in assumption that the rotor optimum inlet flow conditions are largely figures 1 and 2. At design pressure rise the re-staggering process similar for the three builds therefore appears reasonable. for the 50% reaction blading attains a maximum to minimum flow

The characteristic curves also indicate how a line through the range of 1-54 with compressor efficiencies over 0.86. For the 80% Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 stall points tends to turn towards the vertical as the stagger is reaction blading the re-staggering process attains a maximum to increased. This trend in the performance of a compressor using minimum flow range of 1.67 with compressor efficiencies over variable geometry has been noted in other studies (Casey [1988], 0.84. Baumann [1962]) although perhaps not to such an extent as is In order to compare the performance from the two bladings the observed in this study. Although there is a clear shift in the stall normalised characteristic curves from all builds are employed. flow coefficient between Builds 1 and 3, this is not the case These are grouped together and are presented in figure 3 where between Builds 1 and 2. This would suggest the premature stalling performance data are normalised by the optimum values of the of Build 2. Stall margin (S.M) for each build is obtained using the design build. This figure is the essence of this study and clearly expression: indicates the gain in operating range obtained with the high reaction blading. This gain is apparent when all builds are considered. It is also interesting to note that operating range S.M. 1. - (9) increases as the degree of reaction is increased even by a re co staggering process. It is most interesting to note how the efficiency curves suggest where 0, is the flow coefficient at stall point and 0,, is the flow efficient operation is achieved even at the higher reaction levels. coefficient at optimum efficiency point In this way the stall This observation reinforces arguments suggesting a weak margins are calculated as 0.12, 0.18 and 0.22 for Builds 2. 1 and dependence of efficiency on stage degree of reaction, at least for 3 respectively. These indicate a clear gain in stall margin as stator the Mach number range studied. stagger is reduced. A possible explanation of improved range with high reaction may An alternative parameter to consider is the Operating Range be as follows. In the rotor blade boundary layers, the velocity (0.R.) which is defined as: relative to the blades is decreased and the whirl velocities rise towards the local blade speed. The radial static pressure gradients are such as to give approximate radial equilibrium of the O.R. c - 1. (10) 4), mainstream whirl velocity but the boundary layers are not in equilibrium but tend to migrate under pressure and centrifugal forces. The centrifugal forces are largely similar for both builds but where tpc is the flow coefficient at 70% efficiency. This is the pressure gradient is lower for the 80% reaction case (less calculated as .446, .651 and .705 for builds 2, I and 3 respectively, whirl). As the pressure gradient tends to oppose the migration due again indicating the increased range achieved with lower stagger. to the centrifugal effects which are generally the larger, the increased reaction will tend to increase migration. In this way the 80% Reaction Blading Tests blade boundary layers tend to be re-energised by the radial The change in optimum flow condition between Builds 4, 5 and displacement and this may delay separation and hence increase the 6 is illustrated using the same analysis used for the 50% reaction range before stall. builds. From Build 4 the optimum rotor inlet flow angle a, is 61 0 The stalling behaviour hysteresis and type of stall of the six builds tested was found to be different. Exact measurements for calculated for 0.4 = 0 and Cx/U = 0.555. For Build 5 with a, = hysteresis have not been carried out, but an indication of this was 61° and a, = 10° the optimum flow coefficient provided by this obtained, for each build, by noting the time required to come out analysis is 05. which is within 3% of the observed value. For of stall while the electrically operated throttle was opened at a Build 6, with a, = 61° and a, = -10° the optimum flow coefficient constant rate. The stalling behaviour of the builds tested are of 0.61 is within 1% of the observed value. summarised as : When the flow coefficient at stall is considered, this only shows 50% Reaction 80% Reaction a small shift as the stators are re-staggered. The surge margins are Stators -10° Abrupt stall, Abrupt stall, calculated using equation (9) as 0.17 for Build 5, 024 for Build 4 Large hysteresis Large hysteresis and 0.30 for Build 6. These values again clearly indicate the Design Comparative Datum Abrupt stall, improvement in surge margin as stagger is reduced. A similar Small hysteresis trend is observed when the range of operation is considered (544, Stators +10r Part-span stall, Part-span stall. .738 and 1.014 for Builds 5,4 and 6 respectively). Small hysteresis Small hysteresis

5 These results tend to agree with the general trends observed by follow the design values. There are also indications of flow others which indicate that high stagger blading offers better underturning for stations above mid-span and flow overturning for resistance to reverse flow and therefore to abrupt stall with large stations below mid-span. These results tend to indicate corner

hysteresis. vortex formation extending well into the free stream region. Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 As the flow is reduced from optimum to stall conditions, a RADIAL TRAVERSE MEASUREMENTS change of incidence on to the stator of around 17° produces In order to evaluate the variation of typical aerodynamic changes in deviation from 10 for stations below mid-span to 2.5° parameters as the mass flow is varied, traverse measurements were for stations above mid-span. This trend is in agreement with most carried out at two flow conditions referred to as near 'optimum' correlations, Creveling • [1969] McKenzie [1990], suggesting a and 'near stall'. The following sections present and, in a qualitative larger change of deviation with incidence for the higher S/C ratios. marmer, discuss the radial distributions of flow angles. At stall flow conditions there is an indication of flow overturning towards I.D. which is peculiar and suggests strong secondary flow 50% Reaction Blading Tests activity. Build 1. Flow angles for this build are indicated in figures 4 and The flow out of the rotor is shown in figure 6. At optimum flow 5. conditions the flow follows the design angle for most of the blade At optimum flow conditions, the outlet flow angle from the stator with evident undertuming from 88% span to OD. and from ID. to is close to the design values up to around 70% blade height and around 28% span. The underturning at OD. is probably associated for the rotor up to around 45% blade height. Examining the with tip leakage flows and the accumulation of low energy fluid. geometry of this blading, it is observed that while the stator During the design of the rotor blades there was some concern for camber is nearly constant from ID to 0.0, rotor camber decreases the blade loading, as indicated by the diffusion factor, at the I.D. with increasing blade height Since the blading has been designed stations. It may be that the underturning at rotor ID. stations is a using Carter's [1950] unmodified rule, it may be said that this rule result of local blade stalling. tends to underpredict the deviation for the lower camber sections As the flow is reduced to stall conditions there is a change in 0 - this is now generally accepted, Pollard and Gostellow [1967]). rotor incidence from around 9 at I.D. sections to around 7° at OD. Some of the underwriting found in the upper half of the rotor sections. This suggests that the I.D. sections move off design blades may be attributed to secondary flow behaviour particularly faster than the rest of the blade. The rotor relative outlet flow in regions approaching the walls. Pouagare [1985] explains how angle suggests that for the change of incidence onto the rotor, the rotors ability to centrifuge low energy fluid towards 0.D. tends deviation angle increases up to 2° for stations over mid-span. The to feed the suction side blade boundary layers along its path increase in underturning at rotor I.D. is evident at this flow increasing deviation. Lakshminarayana [1970] pointed out how tip condition, again suggesting flow separation and possible stall leakage flows tend to produce undertuming due to fluid migrating initiation here. from the pressure to the suction side over the blades. Flow underturning is certainly observed at rotor OD. and is indicated to Discussion on restaggering occur at stator I.D. When the stator outlet flow angle was considered, blade deviation There seems to be a small region of stator outlet flow overturning angle was found to have a small dependence on stagger angle around 12% of blade height. This could be an indication of vortex setting. The effect of closing the stators (design +10°) is to reduce activity at stator ID. but no conclusions may be drawn just from the mass flow and degree of stage reaction. At this lower reaction the analysis of this study. Such trends in outlet flow angle are level the stator blades have more camber and the rotor blades less reported in other studies Roberts et al [1985]. Joslyn and Thing than required. Stator blades, therefore, operate with less incidence [1984]. and rotor blades operate with more incidence. Closing the stators As the flow is reduced towards stall the rotor incidence tends to effects the work input to the stage and there was an indication of indicate that the ID. rotor stations move off design faster than that the 0.0. sections producing more work even at optimum flow at the O.D. For the range of incidence experienced by the blades conditions. some 1.50 change of deviation is observed for the stator and some The effect of opening the stators (design - 10°) is to increase the 2.0° for the rotor. This, however, seems to affect only the outer mass flow and level of reaction. The stator blades now have less 65% of blade height for both bladings. The decrease in deviation camber and the rotors more than required. A decrease in rotor observed at stator ID. sections as the flow is reduced towards stall, blade incidence and an increase in stator blade incidence was may also be attributed to secondary flow behaviour Joslyn and noted. There was a tendency for the LD. sections to produce more Dring [1984]. work which is in contrast to the effect brought about by the closing of the stator blades. 80% Reaction Blading Tests Build 4. The flow angle out of the stator is shown in figure 7 CONCLUSIONS where, at optimum flow around blade mid-span, the flow angles Contrary to earlier suggestions implying lower efficiencies for

6 high reaction binding this study has demonstrated that well TurboicOmpressoren in Industriellen, Einsaa 8/9, Essen, Nov. 1988 designed high reaction blading can give improved stall margin at Creveling H.F., "Axial-Flow Compressor Computer Program for no loss of efficiency. In fact, the findings of this study indicated Calculating Off-Design Performance. NASA CR 72472, 1969 a slight improvement in optimum efficiency for the 80% reaction Cumpsty N.A., "Compressor Aerodynamics", Longman Group blading. Considering, however, the different methods employed U.K. Ltd., 1989

for the design of the two bladings it might be suggested that such Davis Wit., Millar D.AJ., "Axial Flow Compressor Analysis Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 an improvement in efficiency is a result of the modem methods using a Matrix Method", Carleton Univ. Report MEJA 73/1. 1973 employed for the new 80% reaction blading design. Dong Y., Gallimore Si, Hodson HP., "Three Dimensional Rows The claims for improved stall margin are clearly indicated by the and Loss Reduction in Axial Compressors", Trans. of ASME J. of experimental findings. A wider range of stable operation was also Turbo. 109, pp 354-361, 1987 noted. Such trends for improved stall margin and stable operating Dring R.P., Joslyn H.D., "Through-Flow Analysis of a Multistage range have been observed even when the increase in level of Compressor, ASME 86-GT-13, 1986 reaction was brought about by a re-staggering process. Farmalcalides C.D., "The Effects of Reaction on Compressor The reduction of rotor aspect ratio will contribute to the increased Performance", PhD. Thesis, Cranfield Institute Technology, 1992 range, but the fact that the axial chord is only marginally increased Gostelow J.P., Horlock J.H., Marsh H., "Resent Developments implies that this is a natural advantage of high reaction. The lower in the Aerodynamic Design of Axial Flow Compressors". aspect ratio of the stsuor is not judged to be significant due to its I.Mech.E. E1833N, April 1969 small contribution ID the stage static pressure rise. Hans-Otto J., Heinrich V., "Axial Compressor as a Main Air High reaction designs are in general believed to be limited by the Blower in FCC Units", Turbomachinery International, April 1984 rotor relative inlet Mach number particularly at the O.D. It might Howell A.R., 'The Aerodynamics of the Gas Turbine", I. Aero. be argued, therefore, that although advantages in operating range Soc.. 52, pp 329-356, 1948 (as indicated by the findings of this study) are possible using high Horlock J.H., "Axial Compressors", Butterworth 1958, reprinted reaction blading its application is limited, in jet engines, when the with supplementary material, Krieger Publishing Co. Inc. 1973 Mach number considerations are important and have to be Howell A.R., "Fluid Dynamics of Axial Compressors", Proc. accommodated. It might also be argued, however, that although I.Mechl. 153, pp 441482, 1945 modem transonic compressor designs often involve high Math Johnsen LA., Bullock R.0, "Aerodynamic Design of Axial Flow numbers in the outer diameter sections of the first stages that such Compressors", NASA SP-36, 1965 designs employ more suitable blade profile shapes (such as double Joslyn D.H., Dring R.P., "Axial Compressor Stator circular arc) and that such profiles, if adopted with high reaction Aerodynamics", ASME paper No 84-GT-99, 1984 designs, may be applicable to high Mach number applications. Lieblein S., Schwenk RC., Broderic RI., "Diffusion Factor for Further, the low rotor camber blade sections found in high reaction Estimating !Asses and Limiting Blade Loadings in Axial Flow designs increase the critical Math number for such blade sections Compressor Blade Elements". NACA RM E531D01, 1953 and this may partly offset the penalties associated with the high Leiblein S., "Incidence and Deviation Angle Correlations for rotor inlet Mach numbers. Compressor Cascades", ASME J. of Basic Eng., pp 575-587, 1960 Lakshminarayana B., "Methods of Predicting the Tip Clearance REFERENCES Effects In Axial Flow Turbomachinery", Journal of Basic AGARD AR - 175. "Throughflow Calculations in Axial Engineering, pp. 467-482, 1970 Turbomachines", October 1981 McKenzie A.B., 'The Design of Axial Compressor Blading Based AGARD LS- 167, "Blading Design of Axial Turbomachines", on Tests of a Low Speed Compressor", Proc. I.MechE. 194, pp May 1989 103-111, 1980 Balla! D.R., "Axial Flow Compressor - Stage Stacking McKenzie A.B., -The Selection of Fan Blade Geometry for Techniques", PhD. Thesis, C.I.T., Sept. 1968 Optimum Efficiency", Proc. I.Mech.E. 202, 1988 Baumann H., Schmidt-Theuner P., "Axial Compressors with McKenzie A.B. - Off-Design Performance Data, Unpublished, Adjustable Stator Blades",Brown Boveri Rev. 62 1.51-154, 1962 Private Communication, 1990 • Calvert J., - Private Communication, April 1988 McNally W.D., "Fortran Program for Calculating Compressible Carter A.D.S., 'The Low Speed Performance of Related Aerofoils Laminar and Turbulent Boundary Layers in Arbitrary Pressure in Cascade", Aero. Res, Council CP 29, 1950 Gradients", NASA TN - D568I, May 1970 Casey M.V., "A Mean Line Prediction Method for Estimating the Pollard D., Gostellow J.P., "Some Experiments at Low Speed on Performance Characteristic of an Axial Compressor Stage", Proc. Compressor Cascades", Jnl. of Eng. for Power, Trans. of ASME I.Mech.E., paper C264/87, 1987 89, 1967 Casey M.V., Hugentobler 0., 'The Prediction of the Performance Pouagare M, Galmes J.M., Lakshminarayana, B., "An of an Axial Compressor Stage with Variable Stator Vane Stagger Experimental Study of the Compressor Rotor Blade Boundary Angles". V.1).1. Conference, Thermische Stromungsmaschinen, Layer", Trans of ASME, Vol. 107, pp 364-373, April 1985 Roberts W.B., Serovy G.K., Sandercock D.M., "Modelling the i r 610 5.6 6.0 6.6 7.0 7.5 8.0 3-D Flow Effects on Deviation Angle for Axial Compressor r (cm) 14.22 15-24 16.76 17.78 19.05 20.32

Middle Stages", ASME 85-GT-189, 1985 41) 54.8 52.7 50.0 48.3 46.4 44.5 Roxbee-Cox H., "British Aircraft Gas Turbines", J. Aero. Sci. 13, a. 341.2 32.4 30.0 28.5 26.9 25.5 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 1946 D, 54.8 52.7 50.0 48.3 46.4 44.5

Saunz J.M., "Automated Design of Controlled Diffusion Blades", P. 27.2 25.1 22.7 20.9 19.6 17.6

ASME paper No 88-GT-139, 1988 a' 0.0 0.0 0.0 0.0 0.0 0.0 Seippel C., "The Development of the Brown Boveri Axial 5' 7.0 7.3 7.3 7.6 7.4 7.6 Compressor",Brown Boveri Rev., May 1940 0' 23.1 27.7 27.4 26.4 27.0 26.6

Serovy G.K., "Recent Progress in Aerodynamic Design of Axial 41.0 38.9 36.4 34.6 32.9 31.2

Flow Compressors in the United States", J. Eng. Power, pp 251 S/C 0.79 0.85 0.93 0.99 1.06 1.13 261, July 1966 C (cm) 105 3.05 105 105 105 3.05

Weber A., Faden M., Starken H., Jawtusch V., "Theoretical and 20.6 20.3 20.0 19.8 19.5 _19.0 , Experimental Analysis of a Compressor Cascade at Supercritical Flow Conditions", ASME 87-GT-256, 1987 50% REACTION STATOR DESIGN PARAMETERS TABLE 1 ACKNOWLEDGEMENT The authors wishes to acknowledge the support of Sulzer Esther Wyss, and in particular Dr M.V.Casey and Dr PDalbert for their Rotor assistance and general advice. r (cm) 14.22 15.41 16.51 17.27 17.54 18.51 19.43 20.32

a, 53.92 56.08 57.88 59.04 59.42 60.76 61.93 62.98 5.6 6.0 6.6 7.0 7.5 8.0 r (in) a, 29.38 36.48 41.84 45.0 46.01 49.34 52.07 54.35 r (cm) 14.22 15.24 16.76 17.78 19.05 20.32 p, 52.12 55.76 58.66 60.46 61.09 63.22 65.18 67.02 a, 0.0 0.0 0.0 0.0 0.0 0.0 D, 23.02 30.19 35.55 38.59 39.65 42.81 45.23 47.09 . a. 34.2 32.4 30.0 28.6 26.8 25.5 5 6.38 6.32 6.34 6.41 6.44 6.63 6.97 7.43 13, 0.0 0.0 0.0 0.0 0.0 0.0 .2.58 Lr 3.80 2.21 1.01 029 0.04 -0.84 -1.69 p. 42.1 39.9 36.9 35.2 33.0 31.4 0 29.10 25.56 23.11 21.87 21.44 20.40 19.95 19.93 0.0 0.0 0.0 0.0 0.0 0.0 37.57 42.98 47.11 49.25 50.37 53.01 55.20 57.06 6' 7.9 7.5 6.9 6.6 6.2 5.9 S/C 0.501 0.574 0.651 0.709 0.731 0.816 0.906 1.002 0' 42.1 39.9 36.9 352 33.0 31.4 o 1.998 1.741 1.536 1.410 1.368 1.226 1.104 0.998 16.5 15.7 4. 21.0 19.9 18.4 17.6 C (cm) 4.698 4.436 4.193 4.026 3.968 3.752 3.547 3.353 S/C 0.79 0.85 0.93 0.99 1.06 1.13

. C (cm) 3.05 3.05 3.05 3.05 3.05 3.05 34.2 32.4 30.0 28.6 26.8 25.5 Stator

IGV r (cm) 14.22 15.41 16.51 17.27 17.54 18.51 19,43 20.32

50% REACTION BLADING DESIGN PARAMETERS a, 38.99 36.77 34.90 33.69 33.29 31.88 30.64 29.54

a. 0.0 0.0 0.0 0.0 0.0 0.0 0.0 00 11 42.35 40.45 38.66 37.44 33.29 31E8 30.64 29.54 .9.16 r (in) 5.6 6.0 6.6 7.0 7.5 8.0 P. .9.85 -9.80 .9.69 -9.60 -9.56 -9.43 .9.28 14.22 15.24 16.76 17.78 19.05 20.32 r (cm) 5 9.89 9.84 9.73 9.60 9.61 9.48 9.33 9.20 39.4 44.2 50.0 53.1 56.4 59.0 a, (.. -0.56 -123 -1.64 .I.83 .1.90 -2.07 .2.16 .2.22 5.80 16.5 30.0 37.0 43.8 48.1 a. 0 52.20 50.25 48.35 47.03 46.57 44.91 43.35 41.92 42.4 46.2 50.0 51.6 53.9 55.5 K 16.23 15.33 14.49 13.92 13.72 13.03 12.39 11.80 7.3 22_8 30.9 39.0 44.1 0: S/C 0.634 0.687 0.736 0.770 0.782 0.825 0.867 0.906 0.0 1.50 2.50 3.50 a 1.377 1.455 1.359 1.299 1.279 1.211 1.154 1,104 9.70 9.20 7.20 6.10 4.80 4.00 5' C (cm) 3.808 3.808 3.808 3.808 3.808 3.808 3.808 3808 i 46.3 38.9 27.2 21.7 14.9 11.4 TABLE 2 : 80% REACTION BLAD1NG • FINAL DESIGN PARAMETERS 19.2 26.7 36.4 41.2 46.5 49.8 1.10 S/C 0.77 0.82 0.50 0.96 1.03 105 3.05 3.05 C (cm) 3.05 105 105 46.3 38.9 27.2 21.7 14.9 11.4

50% REACTION ROTOR DESIGN PARAMETERS

1.0

0.95

0.9

LI 0.85 5 LI u.. a. 0.8

0.35 0.75

1.2 Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 0.3 .7 1.1 1.0

0.25 NORMALISED

i) 0.9 l 0.2

Ma 0.8 / 0.7 0.15 -5

0.6 DES 0.1 •10° IAPs 0. 0.4 0.05 0.3

0 02 0.4 0.45 0.5 0.55 0.6 0.65 0.7 0.75 0.8 0.85 0.8 0.9 to 1.1 1.2 1.3 1.4 1.5 1.6 1.7 MASS FLOW COEFFICIENT C y/ U m ICx/UrnIrel

FIGURE 1. STAGE AVERAGED CHARACTERISTICS FIGURE 3. STAGE AVERAGED NORMALISED CHARACTERISTICS WITH VARIATION OF STATOR STAGGER 50% & 80% R. &AGING 50% - REACTION SLAVING

0.95 65- .9 INLET ANGLES FLOI .0-0" • • 60- Sigkle'r.1.49. N. 5g.cr I"... .85 0 55- .8 .... .Cf C. '00..0.0 •••• .7s T.• Sal- ADE i.• ii, • 41.4 '4 0.7 z \ Ow - 45- .te 0.65 E 4 a et 40- ' .. 49% 4 .& 0.6 .:14c p..ti..S. % 0.55 K2/ 35 4*Y".A''jev , 0.5 0.5 d 30- at, 0.45 0.45 a( , 25 .e.23 OUTLET ANGLES > 0.4 0.4 at t;1 20- 0.35 d cc = 0.3 15 - . 0.25 10

er 0.2 5 41 0.15 0 0.1 0.05 0. -10 0.4 0.45 0.5 0.55 0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 10 20 30 410 50 60 70 80 90 100 MASS FLOW COEFFICIENT CxlUm V. GLADE HEIGHT

FIGURE 2. STAGE AVERAGED CHARACTERISTICS FIGURE 4, ROTOR INLET E. OUTLET AIR ANGLES WITH VARIATION OF STATOR STAGGER SOY. - REACTION BLADING BO% - REACTION BLADING STATORS AT DESIGN STAGGER ANGLE SO BO 7 75 70 70 65 65 • •. 60 • 60 55 - 46. 46..on.o..03407.LLL - O• t.7) 55 - FLOW — SO uJ 46, INLET v-ao-oo-o-o L.% 50 ccw 45 tad ANGLES Downloaded from http://asmedigitalcollection.asme.org/GT/proceedings-pdf/GT1994/78835/V001T01A143/2404462/v001t01a143-94-gt-456.pdf by guest on 02 October 2021 wo 40 — 45 BL ADE w 35 • --. DESIGN 40 g 2 3 OPTicit"3C-0.0-0-gMUM FLow too..00.4$70- 41;t- - • -. ."7 3 2$ 2o 8 15 10 OUTLET goW ANGLES +- 0 ' StetirS DESIGN -5 oftroiun FLOW BLADE 10 -15 0 0 10 20 30 40 50 60 70 80 93 100 -20 1111 0 10 20 30 40 50 60 70 80 90 100 % BLADE HEIGHT V. BLADE HEIGHT FIGURE 5. STATOR INLET L OUTLET AIR ANGLES 50% - REACTION FILADING FIGURE 7. STATOR INLET & OUTLET AIR ANGLES STATORS SET AT DESIGN STAGGER ANGLE 80% - REACTION BLADING STATORS AT OESIGN STAGGER ANGLE

75 -

70r

4o. oo. STALL FLOW cec• 65 -0•0•43.0..0.0.• INLET 2.74 w 60 \ ANGLES cgo

OPTIMUM FLOW

ANGLES r4461

25-

20 „ 10 210 30 40 SO 60 70 80 90 100 % BLADE HEIGHT FIGURE 6. ROTOR INLET 8 OUTLET ANGLES BO% - REACTION FILADING phdrovi STATORS AT DESIGN STAGGER ANGLE