Poor Blade Curvature - A Contributor in the Loss of Performance in the Compressor Unit of Gas Turbine Systems

Transcription

1 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 Poor Blade Curvature - A Contributor in the Loss of Performance in the Comressor Unit of Gas Turbine Systems Chigbo A. Mgbemene Deartment of Mechanical Engineering, University of Nigeria, Nsukka 411 chigbo.mgbemene@gmail,com, Abstract The relationshi between the curvatures of the blade and the loss of erformance of the comressor unit of the gas turbine system was studied. Three blade sets of three different blade curvatures o, 35 o and 5 o were fabricated for this investigation. The blades were tested in a manually fabricated wooden wind tunnel and oints of flow searation and vortex height on the trailing edge of each blade set were recorded. The obtained results were then analyzed with resect to the blade velocity distribution effects, comressibility effects and blade loading effects. The analysis indicates that the larger the blade curvature the higher the ossibility of oor erformance of the system. Keywords: blade curvature, flow searation, vortex height, velocity, blade loading, stagnation, ressure, suction surface. 1. Introduction The rofile of the gas turbine blades, be it the comressor or turbine blades, has roven to be the most sensitive area of study in the gas turbine lant. Highly exensive brainower and a lot of man-hours are invested in order to find the blade configurations which will be stable and efficient within secified working angles. Blades must be designed to have correct aerodynamic shae and also be light, tough, and not rone to excessive noise and excessive vibrations. Blades must also be designed to achieve substantial ressure differentials er stage. All these mentioned have direct or indirect relationshis to the rofile of the blade. In actual situation, the rofile of the comressor and the turbine differ. Higher ressure differentials er stage are achieved in the turbine unit than in the comressor hence fewer stages of turbines are required to drive the comressor unit. The turbine blade rofile exhibits deeer curvatures than that of the comressor. This is because the air flow in the comressor is more sensitive to the curvature of the blade rofile than it is in the turbine blade rofile and if dee curvatures are attemted, the ensuing effects are generally undesirable. Studies have shown that if such curvature is attemted in the comressor blade rofile, air flow will tend to searate from the blade surface leading to turbulence, reduced ressure rise, stalling of the comressor with a concurrent loss of engine ower and lowered efficiency of the system [1], all which could be termed loss of erformance. Theoretically, deeer curvatures should give higher system erformance but in the actual situation, loss of erformance of the system occurs. For examle, stalling of the comressor airfoil blade (which is a loss of erformance arameter) results when flow searation occurs over a major ortion of the airfoil s suction surface. If the angle of attack of an airfoil is increased, the stagnation oint moves back along the ressure surface of the airfoil. The flow on the suction surface then must accelerate sharly to round the nose of the airfoil. The minimum ressure becomes lower, and it moves forward on the suction surface. A severe adverse ressure gradient aears following the oint of minimum ressure; finally, it causes the flow to searate comletely from the suction surface inducing form drag and the airfoil stalls [, 3]. A detailed resentation of the review of the studies on stalling and surging of the comressor could be found in Ref. [4]. There are other influences in the system which lead to the oor erformance for examle, according to Cardamone in Ref. [5] the state of the boundary layer influences, in a major way, the loss develoment and this loss develoment directly bears on oor erformance of the system. You and Moin in Ref. [6] resented that flow searation on an airfoil surface is related to the aerodynamic design of the airfoil rofile; and that the nature of flow searation on an airfoil is closely related to the erformance of the system. From the resentations of Refs. [5, 6] and Hulse et al [7] flow boundary layer searation is related to the 6

2 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 airfoil rofile and the erformance of the system. Ref. [7] and Bertagnolio et al [8] were able to relate that the blade tye affects sound generation by the blade while showing that noise level is also a form of erformance measure of the system. They resented that the variations of the searation oint along fixed blades, in turn, roduce changes in the form of their wake and contribute to the discrete-frequency interaction noise from the rotor. Indeed, several studies have been carried out on the erformance of the axial comressor blade but the studies have concentrated more on the stalling and surging of the axial comressor. But again it has been stated that careful design of comressor blading is necessary to revent wasteful losses and minimize stalling [9]. To the ercetion of the author, the effect of the curvature, even to its ossible contribution to the stalling and surging of the comressor has not been shown in simle terms for easier understanding. This write-u therefore aims at resenting the relationshi between the curvatures of the blade and the loss of erformance of the gas turbine system. The effect of the curvature with resect to this loss of system erformance was surmised in this aer from the testing of blades with different curvatures. This was done by investigating the relationshi between curvatures of the blade with resect to the oint of flow searation and vortex height on the trailing edge and subsequently relating them to the velocity distribution effects and blade loading effects due to comressibility effects on the axial tye system. This study was carried out by fabrication of model blade cascades at three different blade angles of o, 35 o and 5 o. Each cascade was tested and the results obtained were analyzed. The aer is meant to give the reader a basic idea of the effects of blade curvature in the loss of erformance of the gas turbine system.. Exerimental Aroach.1 Evaluation of the blade rofile According to Cohen [1] two major requirements of a blade row, whether rotor and stator, are to turn the air through the required angle (β 1 β ) in case of the rotor and (α α 3 ) in the case of the stator; and secondly, carry out diffusing rocess with otimum efficiency i.e. with minimum loss of stagnation ressure. Consequently, the angle at which the air flows across the blades is critical to the erformance of the comressor [11]. One fact remains that air will not leave a blade recisely in the direction indicated by the blade outlet angle, factors like velocity of the air and ressure ratio will affect the exit oint. For this reason, to obtain a good erformance over a range of oerating conditions, it is wise not to make the blade angle equal to the design value of the relative air angle. According to Cohen [1] the design of blades is much of an art. The design of the blade starts with the sketch of the base rofile like as in Fig. 1, from which the blade shae is obtained. The base rofile can be constructed or can be determined from the Joukowski transformation of a given circle. It can be defined as the basic shae of an airfoil with a straight camber line (where the camber length equals the chord) as shown in Fig. (a) [1]. For subsonic flows, it is necessary to use airfoil section blading to obtain a high efficiency but for increased Mach number flows, blade sections based on arabolas are more effective [1]. For this work, the exeriment was carried out in subsonic flow therefore; the base rofile with the airfoil shae was used for the blade design. The rotor and stator blades have the same rofile and camber arc but are arranged as shown in Fig. (b). The base rofile generally used for gas turbine blade shae design is the NACA 4-digit series [1]. Details about the design of blades abound in literatures such as in Refs. [1, 1, 13, and 14] but that is not the oint of this aer; therefore, this will not be fully discussed.. The blade design The ordinates of the base rofile t 1 and t are given at definite ositions along the camber-line as shown in Fig. (a). With the ordinates known the value of the Y U and Y L are comuted as Y, Y U L tn l = (1) 1 where t n = already secified ordinates in Fig. (a) as t 1 and t n = 1 and for uer and lower sections of the rofile resectively l = camber length (chord). 7

3 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 According to Perkins and Hage [1], in the design of the blade, the centre of the leading edge radius is usually obtained by drawing a line through the end of the chord with sloe equal to sloe of camber line at.5 ercent of chord and laying off the leading edge radius along this line. The osition of maximum thickness is always 3 ercent of the chord [1, 15]. The ratio of maximum thickness and chord (t/c) for blade shaes is generally 1 ercent [16]. Following this, the Y values (Y U and Y L ) were comuted and laid off at calculated oints on the X axis. A chord length of 7mm and a blade height of 1mm were chosen for the sake of handling and the size of the wind tunnel that will be used for the testing. A suitable itch/chord ratio (s/c) of.57 (to achieve a high air deflection) [1] and a stagger angle of 34 o were chosen. The choices were made based on the actual data from blade samles obtained from gas turbine blades of Afam Power Station Port Harcourt Nigeria which were Asea Brown Boveri gas turbines whose deflection curves agreed with the tyical design deflection curve resented in Ref. [1]. Since the blades were to be curved, a circular arc camber line was assumed and the blades were constructed for θ = o, 35 o and 5 o. The blades oerate in cascades with notation as shown in Fig. 3. Figure 4 shows the constructed blades (three sets) which were used for the subsequent tests..3 The low seed wind tunnel (smoke tunnel) The exeriment warranted setting u a secial test rig to test the designed blades. A secial wind tunnel - an oen circuit (or Eiffel) tye tunnel was constructed. Because of the size and nature of the exeriment, a smoke tunnel was found suitable for the tests [17]. Smoke was introduced ahead of the model being tested. Very low flow seeds were emloyed for the test to roduce a laminar flow into the test section and to avoid the diffusion of the smoke. The wind tunnel was made of wood. It consisted of a smoke ot, a circular duct, rectangular lenum, a rectangular duct with adjustable section (inflow duct), transarent Persex to, an exit duct, a rectangular chimney and a suction fan. Excet the suction fan, every other equiment was locally fabricated with either wood or aer. The arrangement of the wind tunnel is as shown in Fig. 5. Figure 6 shows front and lan view of the constructed smoke tunnel with a set of blades being tested. The smoke was generated by burning sent engine oil (SAE 4) on embers in a smoke ot and was channeled into the inflow duct through a circular duct. The size of the inflow duct of the tunnel was adjusted such that air (smoke) entered each set of blades at its designed β 1 angle. The air (smoke) was drawn in through suction by the fan laced at the end of the chimney. The fan seeds were measured with a Kestrel 3 digital meter caable of measuring air velocity, temerature, and humidity. Provision was not made for the variation of the sacing between the blades. The observation of the behaviour of air through the blades was visual hence the use of smoke as the fluid medium. 3. Methodology The tests were carried out with the smoke moving at subsonic seed. Different low seeds were used according to the fan suction seed settings. The seeds were: (i) (ii) (iii) N. seed (NS) 3.3m/s Seed 1 (S1) 5.m/s Seed (S) 1.m/s Each blade set was laced between the adjustable duct and the exit chimney. The fan (at a given seed) sucked the smoke through the set of blades. The behaviour of the smoke through the set of blades was observed through the Persex cover. The oints of searation from the leading edge of the blade were recorded as well as the height of the vortex formed on the suction surface at the trailing edge. These were done for the three different fan seeds at three different incidence angles (denoted as in Fig 7) and the results recorded. The average values of each set were obtained and were used in the subsequent lot. Although four different sets of lot namely: (i) (ii) (iii) (iv) Vortex vs searation grah Variation of angle vs average searation grah Seed vs average searation grah Average vortex vs seed grah (for the (-1 o ) result only) 8

4 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 could be used to describe the results of the exeriment, just one set vortex vs searation grah (Fig. 8) is enough for exlanations, therefore only that was shown in this aer. 4. Results and Discussions Searation of flow occurs when the boundary layer has traveled far enough against an adverse ressure gradient such that its energy level dros to a level that the boundary layer seed relative to the object falls almost to zero, the flow then detaches from the object and leaves in form of vortices and eddies. From this study, the oint of this searation was found to be affected by, amongst other arameters, the velocity of the flow. The values of the distance of the oints of searation from the leading edge of the blade and the heights of the vortex formed on the suction surface at the trailing edge were recorded and lotted as shown in the Figs Comaring the grahs in sets for the secified angles of o, 35 o and 5 (i = o ); then o, 35 o and 5 (i = +1 o ) and o, 35 o and 5 (i = -1 o ); for the vortex height vs searation oint grah (Fig. 8), it could be seen that as the camber angle was decreased or increased, the oint at which searation started on the blade decreased or increased resectively. The vortex height also increased or decreased along with the camber angle s increment or decrement resectively. For each angle series for examle, 5 (i = o, +1 o and -1 o ), it could be observed that a variation in the angle of attack of the air on the blade (Fig.9) affected the behaviour of the air in the blade set. A general trend was observed that as the angle of attack varied from o, the oint of searation got closer to the leading edge (Fig.1). It was also observed that the negative series (-1 o ) gave lower searation oints but also gave the lowest vortex heights (Figs. 8 and 9), but for the (-1 o ) the air searated on hitting the leading edge of the blade resulting in higher vortex than would be exected given the foregoing trend. A study of the effect of air seed variations with resect to each blade showed that the lower the seed, the higher the searation oint or higher the seed, the lower the oint of searation. In the 5 (-1 o ), a reversal of trend was noticed; instead of searation oint decreasing with increasing seed, it rather increased with increasing seed (Fig. 9a). The searation oint desite this reversal was still lower than that of the secified angle 5 ( o ). This may have arisen due to the air hitting the blade at an angle such that the laminar boundary layer searated. According to Cardamone [5] and Anderson et al [18] in the resence of adverse ressure gradients, if the laminar boundary layer searates, transition may take lace in a shear layer over the searation bubble. Since a turbulent shear layer has much higher diffusion caability than a laminar one, the flow reattaches usually shortly after transition. This may have been the reason for the trend in the 5 (-1 o ). It was also observed that for any given seed the flow stayed longer on the blade suction surface before searation on the shallower curved blades (Figs. 9 and 1). Vortex heights increased as the seeds were increased leading to turbulent wakes. Aart from the abnormal behaviour of the air in the 5 (-1 o ) blade osition, it was observed that generally as the seed was increased the air gradually lost its ability to negotiate curves and then searated from the blade. The degree of searation and formation of vortex deended on the degree of curvature of the blade and also the angle of attack. This effect could be exlained by the work of Perkins and Hage in Ref. [1]. They resented that the degree of searation on an aerodynamic body is largely deendent on the magnitude of the unfavourable ressure gradient to the rear of the oint of minimum ressure or maximum surface velocity. Therefore, if the ressure gradient, d/dx, along the surface from this oint is equal to or less than zero, then no searation exists. But when the ressure gradient is gradual, the searation occurs so near the rear of the body such that only a very small turbulent wake is roduced and the boundary layer is extremely thin everywhere. The drag created by such a body is small and arises mostly from skin friction. If the ressure gradient is high, searation occurs well forward of the rear stagnation oint and a turbulent wake exists which alters the otential-flow icture and ressure distribution. 4.1 The consequences of the curvatures The consequences of a large blade curvature could be seen when the stagnation ressure ratio in a blade row given as Eq. () [1] is looked at. 9

5 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 ( γ 1) γ ( γ 1) s M ω = 1+ [( rv) ( rv) 1] 1 () 1 γ s 1+ M w where s1 and s are the stagnation ressures at stations 1 (ahead) and (aft) of the cascade. From the equation, it could be seen that achieving a large stagnation ressure in the cascade can only be done by having a large curvature but it introduces two deleterious effects velocity distribution and blade loading effects. A large stagnation ressure will also introduce a large adverse static ressure gradient which leads to boundary layer searation on the blade The velocity distribution effect due to the blade curvature Since the air assing over an airfoil will accelerate to a higher velocity on the convex surface and in a stationary row, this will give rise to a dro in static ressure. This convex surface is termed the suction side of the blade. The concave side will exerience a deceleration of the flow and a rise in static ressure hence it is termed the ressure side (Fig 7). Given similar conditions, the velocity distribution through the assage between the ressure side and the suction side will be as resented in Refs. [1, 19] as shown in Fig.11. The maximum velocity on the suction surface will occur at around 1 15 ercent of the chord from the leading edge after which it falls steadily until the outlet velocity is reached. It was found that relatively thick surface boundary layers resulting in high losses occur in regions where raid changes of velocity occur, i.e. in regions of high velocity gradient. From Fig. 7, this would most likely occur on the suction surface of the blade. Then increasing the convex nature of the suction surface would increase the velocity gradient further leading to a serious loss due to friction and breakdown of flow with its attendant dro in stagnation ressure Blade loading effects A large blade curvature will lead to a high blade aerodynamic loading and low values of minimum static ressure on the blade suction side [1] and comressibility effects affect the blade loading in that they change the ressure distributions along the blade so that the blade loadings may become excessive. This ressure distribution is related to the ermissible deflection in the blade. For a given eriheral seed, the ressure rise obtained in a stage is a function of a coefficient of deflection τ, which for a articular degree of reaction deends on the ermissible deflection β of the flow in the cascades []. This was shown by Carter [1] and Zweifel []. Let us consider the study done by Carter based on observations of an N.G.T.E.[3] rofile test data as shown in Fig.1, measured in a 5 in low seed cascade tunnel of NACA for subsonic flow; here θ = β = 18.6 o, the static ressures on the rofile was obtained by means of the ressure coefficients c. The values of c at stations (1) and () are given as: c c s1 1 1 = (3) ( ρ / ) V 1 s = (4) ( ρ / ) V where s1 and s are the total stagnation ressures ahead and aft of the cascade velocities V 1 and V resectively. For an incomressible flow with friction, it is ossible to exress the frictional losses by the dro of the absolute stagnation ressures at the stations ahead (1) and aft () of the cascade. This reduction in absolute stagnation ressure is identical with that of the relative total ressure SR, since the stream surfaces are coaxial cylinders. The loss through a cascade has also been related to the ideal velocity head which could be roduced by a frictionless rocess between the stagnation ressure s1 and the static ressure. However, if the velocity diagram Fig.13 [4] is to reresent the actual velocities that occur for a rocess with friction, then the ressure coefficient c will be given as c( ξ ) SR1 SR S1 S = = (5) ( ρ / ) V ( ρ / ) V 3

6 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 Neglecting geootential energy differences, the difference in relative ressures can be given as SR1 SR = 1 + ( / )( V V ) ρ (6) Combining Eqs. (5) and (6) we have ( ρ / )( V V ) ξ( ρ / ) V 1 = 1 1 (7) The normal force F on the blade can then be reresented as Cos β Sinβ F = sρv Vu ξs( ρ / ) V (8) Cos β Neglecting the second term we have that V u F sρ V (9) Introducing the solidity σ = blade chord/blade sacing ( c / s) σ =, (1) and if we exress the normal force in terms of the lift coefficient C L which is dimensionless, then F =. (11) ρ C L / ( ) V c Combining Eqs. (9) and (11) then C L V = σ V u (1) From Fig.13 C L can then be exressed as C = [ tanβ1 tanβ ] cosβ (13) L σ From Eq. (13) the link between the ermissible deflection and the ressure distribution could be deduced via the lift coefficient. It could also be deduced that the force on the blade deends on the ressure gradient on the suction side of the blade which in turn is affected by the degree of fluid deflection. Therefore, increasing the blade curvature ends u increasing the blade loading. 5. Conclusions The relationshi between the curvatures of the blade and the loss of erformance of the gas turbine system has been studied. The discussions have shown that as the curvature was increased from o to 5 o, the oints of searations receded and vortex heights increased and consequently, the detrimental factors that lead to stall and surge in the system were accentuated. Blade aerodynamic loading and the velocity gradient increased with increasing curvature leading to a serious loss due to friction and breakdown of flow with its attendant dro in stagnation ressure. Following these, it can then be concluded that large blade curvature in comressors lead to early flow searations with its adverse effects in the whole comressor system. Therefore, designing blades with large curvatures will lead to loss of erformance in the comressor system. Acknowledgements The author would like to thank Prof. H. I. Hart of Rivers State University of Technology Port Harcourt and the management of Afam Power Station, Port Harcourt Nigeria for their suort towards this roject. References [1] Oates, G. C., 1988, Aerothermodynamics of Gas Turbine and Rocket Proulsion (Revised and Enlarged), AIAA Educational Series, American Institute of Aeronautics and Astronautics, Inc., Washington D.C. 31

7 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 [] Kay, J. M. and Nedderman, R. M., 1974, An Introduction to Fluid Mechanics and Heat Transfer, 3 rd Ed., Cambridge University Press, Cambridge, UK. [3] Fox, R. W. and McDonald, A. T., 1978, Introduction to Fluid Mechanics, nd Ed., John Wiley and Sons, New York. [4] Mgbemene, C. A.,, The Effect of Blade Curvature in the Failure of Gas Turbine Blades, M.Eng. Thesis, University of Nigeria, Nsukka, Enugu State. [5] Cardamone, P., 6, Aerodynamic Otimization of Highly Loaded Turbine Cascade Blades for Heavy Duty Gas Turbine Alications, Ph.D Thesis, Universität der Bundeswehr München, Germany. [6] You, D. and Moin, P., 6, Large-Eddy Simulation of Flow Searation over an Airfoil with Synthetic Jet Control, Center for Turbulence Research Annual Research Briefs, , htt://ctr.stanford.edu/resbriefs6/6_you1.df, accessed 16 Jan. 1. [7] Hulse, B. T., Lange, J. B., Bateman, D. A. and Change, S.C., 1966, Some Effects of Blade Characteristics on Comressor Noise Level, FA65WA-163, Federal Aviation Administration, Washington, D.C. [8] Bertagnolio, F., Madsen, H. A. and Bak, C., 1, Trailing edge noise model validation and alication to airfoil otimization, Journal of Solar Energy Engineering, 13, [9] Bhushan, N., 8, Flow through Centrifugal & Axial Flow Comressors, htt:// accessed 3 Feb. 1. [1] Cohen, H., Rogers, G. F. C. and Saravanamuttoo, H. I. H., 1988, Gas Turbine Theory, 3 rd Ed., Longman Scientific and Technical, Harlow Essex, UK, Cha. 4. [11] Anonymous, 1, Fundamentals of Gas Turbine Engines, htt:// accessed 16 Jan. 1 [1] Perkins, C. D. and Hage, R. E., 1958, Airlane Performance Stability and Control, John Wiley & Sons, Inc., New York, Cha.. [13] Dixon, S. L., 1998, Fluid Mechanics and Thermodynamics of Turbomachinery, 4 th Ed., Butterworth- Heinemann, Boston, MA. [14] Horlock, J. H., 1958, Axial Flow Comressors, Butterworths Scientific Publications, London. [15] Carter, A. D. S., Turner, R. C., Sarkes, D. W. and Burrows, R. A., 196, The Design and Testing of an Axial-Flow Comressor having Different Blade Profiles in Each Stage, Reorts and Memoranda, Ministry of Aviation, Aeronautical Research Council, London, A.R.C. Technical Reort No [16] Boyce, Meherwan P.,, Gas Turbine Engineering Handbook, nd Ed., Gulf Professional Publishing, Boston, Cha. 7 [17] Poe, A. and Harer, J. J., 1966, Low Seed Wind Tunnel Testing, John Wiley and Sons Inc., New York. [18] Anderson, D. A., Tannehill, J. C. and Pletcher, R. H., 1984, Comutational Fluid Mechanics and Heat Transfer, Hemishere Publishing Cororation, New York, NY, [19] Elmstrom, M.E., 4, Numerical Prediction of the Imact of Non-Uniform Leading Edge Coatings on the Aerodynamic Performance of Comressor Airfoils, Master s Thesis, Naval Postgraduate School, Monterey, CA. [] Sinette, J. T. Jr., Schey, O. W. and King, J. A., 1943, Performance of NACA Eight Stage Axial Flow Comressor Designed on the Basis of Airfoil Theory, NACA Reort 758, Washington D.C. [1] Carter, A. D. S., 1955, The Axial Comressor, Gas Turbine Princiles and Practice, Sir H. R. Cox (Ed.), G. Newnes Ltd, London, Section 5. [] Zweifel, O., 1945, Otimum Blade Pitch for Turbo-Machines with Secial Reference to Blades of Great Curvature, Brown Boveri Review, 3. [3] Felix, A. R. and Emery, J. C., 1957, A Comarison of Tyical National Gas Turbine Establishment and NACA Axial Flow Comressor Blade Sections in Cascade at Low Seed, NACA Technical Note 3937, Washington D.C. [4] Vavra, M. H., 196, Aero-Thermodynamics and Flow in Turbomachines, John Wiley & Sons, Inc., New York, NY. 3

8 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 The Figures Y Y U X l = c Fig. 1: An aerofoil blade rofile. t Y U L.E. t Y L Camber-line length l Camber line C B D ζ β O β 1 A Chord Centre of camber Rotor Stator Uer Surface Y U (%l) T.E. (a) Lower Surface Y L (%l) L.E. = Leading edge T.E. = Trailing edge O (b) Fig. : (a) The blade base rofile. (b) The rotor and stator layout in blade design. 33

9 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 Rotor c Position of maximum thickness θ = o, 35 o and 5 o β Direction of drum rotation ζ Front of comressor s V δ β 1 β Air flow direction a β 1 Point of maximum camber Radius of camber arc Fig. 3: Comressor cascade and blade notation. i V 1 β 1 = blade inlet angle β = blade outlet angle θ = blade camber angle = β 1 - β ζ = setting or stagger angle s = itch ε = air deflection = β 1 - β β 1 = air inlet angle β = air outlet angle V 1 = air inlet velocity V = air outlet velocity i = incidence angle = β 1 - β 1 δ = deviation angle = β - β c = chord t = maximum thickness a = distance of maximum camber from L.E. (a) (b) (c) Fig. 4: The fabricated blade samles (a) θ = o (b) θ = 35 o and (c) θ = 5 o. 34

10 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 Air (smoke) inlet Inflow duct Blade set under test Suction fan Wind tunnel (fixed section) Inflow duct adjuster Fig. 5: A schematic diagram of the wind tunnel with a blade set in lace. Suction fan Exit chimney (a) Smoke ot Flow control section Blades under test Fig. 6: (a) Front view and (b) Plan view of the wind tunnel. (b) 35

11 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 Suction surface y Air (smoke) in β 1 θ = o, 35 o and 5 o -1 o o Air (smoke) out +1 o x Pressure surface Incidence angle Fig. 7: The blade test nomenclature. Vortex height [mm] NS ( 5 o ) S1 ( 5 o ) S ( 5 o ) NS ( 35 o ) S1 ( 35 o ) S ( 35 o ) NS ( o ) S1 ( o ) S ( o ) Point of Searation [mm] Fig. 8: A lot of Point of Searation against Vortex Height for the (a) θ = o, (b) θ = 35 o and (c) θ = 5 o blade sets at i = o. 36

12 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 (a) Vortex height [mm] o o o +1 o +1 o +1 o -1 o -1 o -1 o 5 o Blade Set (b) Vortex Height [mm] o o o +1 o +1 o +1 o -1 o -1 o -1 o 35 o Blade Set Point of Searation [mm] Point of Searation [mm] (c) Vortex Height [mm] 1 5 o o o +1 o +1 o +1 o o Blade Set Point of Searation [mm] Fig. 9: A lot of Vortex Height vs. Point of Searation. 37

13 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 Air in i = o P S P S1 P NS S S1 x (mm NS y (mm) 15 (a) P S P S1 PNS y (mm) Air in i = o S S1 NS x (mm) (b) 44.5 Air in i = o (c) P P S1 S P NS S S NS x (mm) y (mm) Fig. 1: Points of searation P NS, S1, S (x axis) and vortex heights (y axis) for the (a) θ = o, (b) θ = 35 o and (c) θ = 5 o blade sets. 15 V max Suction surface V 1 Mean velocity in Pressure V Chord ercent 1 Fig. 11: Tyical velocity distribution through the assage of a cascade set [1, 19]. 38

14 ISSN 4-33 (Paer) ISSN (Online) Vol., No.4, 1 1 β = 18.6 o Suction side u β 1 = 6 o 15.6 o V o x c l Pressure side V β = 41.4 o c l = l ( ρ / ) V C Pressure side c l A c u = u ( ρ / ) V -1 ⅓ K = Suction side c u -3 B x/c.8 1. Fig. 1: Measured ressure distribution on rofile NGTE 1C4/3C5 in cascade [3]. D F a β 9 o F ε β F e 9 o F u β β 1 V a Axial direction V 1 V V β ΔV V u V u1 V + V u1 u ½ΔV u ½ΔV u Fig. 13: Velocity diagram of axial flow comressor cascade [4]. 39

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