Experimental and numerical investigation of secondary flow on compressor blades

Transcription

1 Experimental and numerical investigation of secondary flow on compressor blades A. Hergt *, R. Meyer * German Aerospace Center (DLR); Institute of Propulsion Technology; Dept. of Turbulence Research; Müller-Breslau-Str. 8; D Berlin; Germany alexander.hergt@dlr.de K. Engel MTU Aero Engines GmbH; Dachauer Str. 665; D Munich Abstract An experimental and numerical investigation on the flow separation in the corner between a wall and a vane in a highly loaded compressor cascade was performed. A large part of the total losses of a compressor stage is caused by this separation. The objective of the investigation is to understand the fluid mechanic mechanism of corner separation and to detect reference results for developing a flow control technique to avoid this flow separation. The experiments with a compressor cascade were carried out at a high-speed test facility at the DLR in Berlin. The experiments were done at Reynolds numbers up to Re = 0.6 x 10-6 (based on 40 mm chord) and Mach numbers up to Ma = 0.7. The profile of this blades represents the 10 % cut of vane length distance from the hub of the guide vanes of the single stage axial compressor of the Technical University of Darmstadt. For the assessment of the total pressure losses of the cascade (caused by the corner separation) a pressure measuring technique was used. To detect the separation area on the vane a flow visualisation technique was used. Different kind of flow control devices are intended to influence the corner separation. In order to optimize the devices for an application in turbomachines, an experimental study is currently ongoing, which investigates the use of such devices to reduce the losses of an aerodynamically highly loaded compressor cascade. In addition to the experiments, numerical computations were carried out with TRACE, a parallel Reynolds-Averaged Navier-Stokes flow solver, which has been developed for the simulation of turbomachinery flow. The computations were carried out with the same geometry as the experiments, including the measured inflow boundary layer conditions at the side walls. Keywords: corner separation, secondary flow, compressor cascade Nomenclature Flow parameter: Geometric parameter: Ma [-] Mach number β 1 [ ] incoming flow angle Re [-] Reynolds number β 2 [ ] out going flow angle p t [Pa] total pressure β S [ ] stagger angle p [Pa] static pressure t [m] pitch q [Pa] dynamic pressure c [m] vane chord ρ [kg/m 3 ] density h [m] vane height ζ [-] total pressure loss coefficient x [m] axial coordinate c f [-] skin friction coefficient u [m] circumferential coordinate TF [-] transition coefficient z [m] vane height coordinate * Scientist, Turbulence Research Group Head, Compressor Aerodynamic 1

2 Introduction A better knowledge about the flow phenomena in turbomachine cascades is necessary to improve the efficiency. The basic for experimental investigations of the phenomena is the two dimensional cascades. This cascade is a simplification as a result of a coaxial cylinder barrel cut of a three dimensional cascade. Since the 1950 s, systematic studies of the flow in two dimensional cascades were performed with the mainly objects of the influence of compressibility, Mach number, Reynold s number and turbulence on cascade flow [7], [9]. Furthermore secondary flow phenomena and cascade losses were investigated. Especially in the 1960 s the attention were turned to profile loss investigation to decrease the total losses of a compressor stage. Additional investigations of secondary flow phenomena were performed in the early 1990 s. [3], [6], [11]. Another major source of losses in a compressor cascade is the separation between the side wall and the vane (i.e. corner separation ) caused by the interference between the wall and vane boundary layer and the high positive pressure gradient in flow direction. The current experiments were performed with the objective of understanding the fluid mechanic mechanism of corner separation and of detecting reference results to permit more investigations in order to improve the efficiency. In addition to the experiments, numeric computations were carried out with the aim of better visualization and understanding of the flow. Due to the small dimension of the test section and the influence of probes in the flow the accessibility for measurements are limited. By means of CFD we obtained more information about flow parameters e.g. the pressure distribution at the vane and the comparison between computed and measured results allowed us to validate the used flow solver also. Setup of the Cascade Experiments The High Speed Cascade Wind Tunnel The experimental investigations were carried out at the high-speed wind tunnel of the DLR (Institute of Propulsion Technology, Department of Turbulence Research) in Berlin (figure 1). The channel has a rectangular cross section of 40 mm width and 90mm height at the exit of the nozzle which has a contraction rate of 1:218. Thus flow velocities of Ma = 0.7 with a Reynolds number of can be obtained [5]. Figure 1: High speed wind tunnel connected to compressor cascade test section 2

3 The test section of the channel with the connected cascade as shown in figure 2 has some special features, which permit the variation of several parameters. Figure 2: Test section with cascade The incoming flow angle β 1 can be adjusted separately from the stagger angle β S. This allows the usage of different blade geometries, with the same cascade. Boundary layer suction at all four channel walls is possible and can be adjusted separately (figure 2). The suction at the upper and bottom walls allows the adjustment of the static pressure over the channel height. Thereby a homogeneous inflow according to an infinite blade cascade is achieved. Another special feature of this particular test-section is the adjustable boundary layer thickness of the side walls. The boundary layer thickness can be reduced by suction at the side walls. To increase the boundary layer thickness the optional spoiler section can be used. Compressor Cascade and Measurement Equipment The cascade used for the investigations consisted of 5 blades with a chord length of 40 mm. The profile of this blades represents the 10 % cut of vane length distance from the hub of the guide vanes of the single stage axial compressor of the Technical University of Darmstadt. The main geometrical parameters and design flow conditions are shown in figure 3 and table 1. Ma 1 = 0.66 c = 40 mm β 1 = 132 t/c = 0.55 β S = h/c = 1 Table 1: Geometrical and design flow conditions Figure 3: The compressor cascade 3

4 The aspe ct ratio (ratio of blade height and chord) is 1. This value is typical for modern highly loaded compressors and was selected in order to let the secondary flow effects dominate. The cascade is mounted in the test section as shown in figure 4. Due to the boundary layer suction at the upper and lower wall (also shown in figure 4) a homogeneous inflow at the center blade (blade 3) of the cascade could be achieved. This was necessary, since the measurement focused on the determination of the total pressure losses of a passage in an infinite cascade. The local total pressure loss coefficient is defined in equation 1 [2],[10]. ζ ( u, z) = p t1 pt 2 ( u, z) q 1 (1) In addition all presented loss results are mass flow averaged. The main inflow conditions and the total pressure distribution behind the cascade must be determined to compute the coefficient. Hence, the static inlet pressure, the total inlet Figure 4: Cross section of the mounted cascade with pressure and the total temperature in the settling reference coordinates system chamber were measured. Furthermore the total pressure distribution was measured with two wake rakes 40% of chord length c behind the trailing edge of the cascade blades, which can be traversed in circumferential direction. One rake consists of 26 total pressure probes in line from one side wall to the other side wall of the cascade. The other consists of one static pressure probe and 4 Conrad probes in order to detect the outflow angle at 4 vane height positions [5]. Thereby it is possible to compute the mass flow averaged local total pressure loss distribution and total pressure loss in vane height direction of a cascade passage. To detect the separation area on vane a flow visualization technique was used. Cascade Losses The total pressure loss coefficient ζ ges of a cascade passage is computed by integration of total pressure loss coefficient distribution ζ(z) in vane height direction, which is composed of three parts of losses, as shown in figure 5 and equation 2: The side wall boundary layer losses ζ BL, the profile losses ζ P caused by the friction of flow around the vane profile with infinite aspect ratio and the losses due to corner separation ζ SP [1], [4], [8]. ζ = ζ + ζ + ζ ges P BL SP (2) Experimental Errors Systematic errors of the experimental setup can be assessed. During the experiments only pressure and temperature were measured. The pressure measurement chain consists of probes, transducer and voltmeter. The total pressure probe error Figure 5: Total pressure loss parts 4

5 caused by probe inflow angle divergence of ±10 amounts to max ±0.5% [13], the transducer error to ±0.05% and the voltmeter error is negligible. Thereby a max measurement chain error of the total pressure downstream of the cascade of 0.55% is possible. Upstream of the cascade the total and static pressure probes error is negligible, since the inflow angle divergence is less than 5. Hence, the total pressure loss coefficient error amount to 0.65%. The temperature measurement chain consists of Pt100 Sensor with an error of ±0.3K and Voltmeter with negligible error. Differences caused by wire length will be eliminated by four wire measurement. Experiment Results The experimental investigation was performed to detect the location of corner separation at the selected vane profile and to measure the value of total pressure losses of a cascade passage. The obtained results serve as reference for developing appropriate flow control devices and for understanding the secondary flow phenomena. In a first series measurements were carried out at the design Mach number of 0.66 and inflow angle at peak efficiency and off design with an incidence of -6.In the presentation an example for positive incidence angle will be also given. The results are shown in figure 6 and 7. No absolute numbers can be given, since the data were obtained in cooperation with MTU Aero Engines. Nevertheless the main information can be obtained from the figures. In figure 6 and 7 the local total pressure loss coefficient distribution ζ(u,z) (mass flow averaged) in the wake, the measured outflow angles and the total pressure loss distribution ζ(z) in vane height direction are shown. The red and magenta areas present high losses and the blue areas less losses. lokal total pressure loss coefficient (mass flow averaged) suction side 0.5 lokal total pressure loss coefficient (mass flow averaged) suction side 0.5 outflow angle [ ] relative cascade pitch u/t [-] 0 pressure side trailing edge outflow angle [ ] relative cascade pitch u/t [-] 0 pressure side total pressure loss coefficient (upstream view) (mass flow averared) relative vane height z/h [-] total pressure loss coefficient (upstream view) (mass flow averared) relative vane height z/h [-] Figure 6: Measurement results at design point (Ma 1 = 0.66) Figure 7: Measurement results at incidence of -6 (Ma 1 = 0.66) 5

6 The high loss areas in the corner between wall and vane (relative cascade pitch u/t = 0 is equivalent to trailing edge position of the vane) in both figures present the losses caused by corner separation. An increase of profile losses and the displacement of corner loss areas due to incidence from design point are identifiable too. The profile loss increase is caused by the visible increase of the wake expansion at the pressure side (figure 7, upper diagram). In the flow visualisation (figure 9, lower picture) a flow separation on the pressure side near the leading edge can be seen, which also explain the loss increase. Figure 8: Flow visualisation on the vane at design point (Ma 1 = 0.66) Figure 9: Flow visualisation on the vane at incidence of -6 (Ma 1 = 0.66) By means of flow visualization (figures 8 and 9) the typical corner separation on the rear part of the suction side of the reference blades can be detected. The displacement of separation areas due to incidence from design point is detectable in the flow visualization figures too. Furthermore a separation bubble on the suction side is shown which give us information about the transition on the vane. Upstream of the bubble the boundary layer of the blade is laminar and downstream it is turbulent. ζ ges [-] total pressure loss Ma = 0.7 total pressure loss Ma = 0.66 total pressure loss Ma = 0.5 In a second series measurements were carried out at varying incidence angles between -8 and +8 and three inflow Mach numbers (Ma 1 = 0.5 / 0.66 / 0.7) to obtain loss curves as shown in figure 10. The peak efficiency of the cascade is identified at the curve which presents a reference for further measuring and development of flow control devices. - incidence + inflow angle β 1 [ ] peak efficiency Figure 10: Measurement results depending on inflow angle for various Mach numbers 6

7 Numerical Investigation The numerical investigations were performed at MTU Aero Engines in Munich. For the numerical simulation a structured OCH grid [12] consisting of five blocks with nodes was generated. The computations were carried out with TRACE, a parallel Reynolds-Averaged Navier-Stokes flow solver, which has been developed for the simulation of turbomachinery flow[14]. leading edge trailing edge incoming flow area of corner separation laminar separation turbulent reattachment Figure 11: Computed static pressure distribution and strike lines on suction side of the vane at design point (Ma 1 = 0.66) Numerical Results The incoming flow parameters for the simulation are based on cascade measurements at the DLR in Berlin. For the stationary computation, which was carried out at design point of the cascade the k-ω turbulence model and the Drela modified Abu Ghannam & Shaw transition model was used [14] and the transition was allowed on the whole blade surface. In figure 11 the computed pressure distribution and strike lines on suction side of the vane is shown whereas the blue area is equivalent to peak suction. The laminar separation bubble on the suction side as indicated in figure 8 is also visible in figure 11. This interpretation is confirmed by the skin friction coefficient c f and the transition coefficient shown in figure 12. At the point of laminar separation the skin friction coefficient c f becomes negative and the transition CF leading edge laminar separation turbulent reattachment X Figure 12: Computed static pressure distribution P, skin friction coefficient c f and transition coefficient TF at mid span (at design point, Ma 1 = 0.66) P CF TF trailing edge P TF 7

8 coefficient TF increases from zero to one. Furthermore the location and area of corner separation is computed and is in good agreement with the experimental results of the flow visualization. During the measurement the mean outflow angle at four vane height positions was determined. The computed and measured outflow angles are shown in figure 13 and coincide very well. outflow angle β 2 [ ] TRACE Computation Measurement The experimental results of the local total pressure loss coefficient distribution (figure 14b) are compared with the computed values (figure 14a). The red and magenta areas present high losses and the blue areas less losses. The position of high loss areas at the measurement result are well reflected in the computed result, only the shape of the high loss area is showing slight differences. This 0 0,1 0,2 0,3 0,4 0,5 relative vane height z/h [-] Figure 13: Comparison of computed and measured outflow angle distribution at design point (Ma 1 = 0.66) can be attributed to the linear eddy viscosity turbulence model since the simulation of the three dimensional turbulent flow in the corner is difficult. Nevertheless, the comparison of computed and measured results (figures 14a and 14b) are in fair agreement. Figure 14a: Computed local total pressure loss coefficient distribution at design point (Ma 1 = 0.66) Figure 14b: Measured local total pressure loss coefficient distribution at design point (Ma 1 = 0.66) Discussion of the Results The experimental results as well as the numerical results allow a good assessment of the fluid mechanic mechanism of corner separation. It is well known that this separation mainly caused by the interaction between the wall and blade boundary layer and the high pressure gradient in flow direction. The low energetic boundary layer flow is decelerated by the increasing pressure till separation. The numerical simulation yields detailed data for the backflow areas in both corners which cannot be obtained easily by experiments. Furthermore the position of separation relative to the pressure distribution is known. The experimental results have also shown that the corner separation losses are an important part of the total losses of a cascade and they will serve as reference for continuing measurements. 8

9 Conclusions E xperimental investigations of secondary flow on compr essor cascade blades were performed. Thereby the values of losses at different incidences and Mach numbers were determined. In addition a n umerical investigation was carried out. The comparison between computed and measured results is showing a very good agreement. The numerical simulation yields more information about the separating flow region and improves the understanding of the investigated flow phenomena substantially. Finally a better understanding of structure and development mechanism of corner losses was achieved. Acknowledgments The investigations reported in this paper were performed within a cooperation project with MTU Aero Engines in Munich. In this manner we would like to thank for the good cooperation. References [1] AMECKE J.: Auswertung von Nachlaufmessungen an ebenen Schaufelgittern, Bericht 67 A 49 AVA Göttingen, 1967 [2] CUMPSTY N.A.: Compressor Aerodynamics, Krieger Publishing Company, 2004 [3] HÜBENER J.: Experimentelle und theoretische Untersuchung der wesentlichen Einflussfaktoren auf die Spalt- und Sekundärströmung in Verdichtergittern, PhD Thesis, Universität der Bundeswehr München, 1996; Germany [4] MEYER R., BECHERT D.W., HAGE W.: Secondary Flow Control on Compressor Blades to improve the performance of axial turbomachines, 5 th European Conference on Turbomachinery Fluid Dynamics and Thermodynamics, Prague, March 2003 [5] MEYER R.: Versuchsaufbau und Auswerteverfahren für ebene Verdichtergitter mit Sekundärströmungsbeeinflussung bei hohen Unterschall- Geschwindigkeiten, DLR-IB /B5, Dept. of Turbulence Research, Institute of Propulsion Technology, DLR, Berlin, 2002 [6] SCHEUGENPFLUG H.: Theoretische und experimentelle Untersuchungen zur Reduzierung der Randzonenverluste hochbelasteter Axialverdichter durch Grenzschichtbeeinflussung, PhD Thesis, Universität der Bundeswehr München, 1990; Germany [7] SCHLICHTING H.: Recent Research on Cascade-Flow Problems, Journal of Basic Engineering, ASME, 1966 [8] SCHOLZ N.: Über den Einfluß der Schaufelhöhe auf die Randverluste in Schaufelgittern, DK : , Rundschau Forschung 20. Band/Heft 5, DF457, Braunschweig,1954 [9] SCHOLZ N.: Über die Durchführung systematischer Messungen an ebenen Schaufelgittern, Zeitschrift für Flugwissenschaften, October 1956 [10] SCHOLZ N.: Aerodynamik der Schaufelgitter, Verlag Braun, 1965 [11] WATZLAWICK R.: Untersuchungen der wesentlichen Einflussfaktoren auf die Sekundärverluste in Verdichter- und Turbinengittern bei Variation des Schaufelseitenverhältnisses, PhD Thesis, Universität der Bundeswehr München, 1991; Germany [12] WEBER A.: 3D Structured Grids for Multistage Turbomachinery Applications based on G3DMESH, DLR IB , Numerical Simulation Group, Institute of Propulsion Technology, DLR, Cologne, February

10 [13] WUEST W.: Strömungsmeßtechnik, Vieweg & Sohn Verlag, Braunschweig, 1969 [14] DLR IB : TRACE USER s MANUAL, Numerical Simulation Group, Institute of Propulsion Technology, DLR, Cologne, March