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-10623 Berlin; Germany e-mail: alexander.hergt@dlr.de

K. Engel‡

MTU Aero Engines GmbH; Dachauer Str. 665; D-80995 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 β1 [°] incoming flow angle Re [-] Reynolds number β2 [°] out going flow angle pt [Pa] total pressure βS [°] stagger angle p [Pa] static pressure t [m] pitch q [Pa] dynamic pressure c [m] vane chord ρ [kg/m3] density h [m] vane height ζ [-] total pressure loss coefficient x [m] axial coordinate cf [-] skin friction coefficient u [m] circumferential coordinate TF [-] transition coefficient z [m] vane height coordinate

* Scientist, Turbulence Research Group ‡ Head, Compressor Aerodynamic

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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 600000 can be obtained [5].

Figure 1: High speed wind tunnel connected to compressor cascade test section

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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 β1 can be adjusted separately from the stagger angle βS. 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.

Ma1 = 0.66 c = 40 mm β1 = 132° t/c = 0.55 βS = 105.2° h/c = 1

Table 1: Geometrical and design flow conditions Figure 3: The compressor cascade

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T ct ratio (ratio of blade height and chord)he aspe is 1. This value is typical for modern highly loaded compressors and was selected in order to

(1)

In add meraged. The main inflow conditions and

e total pressure distribution behind the cascade

ling edge of th be traversed in otal pressure probes in line from one side wall to the other

efficient ζges of a cascade by integration of total

(2)

mental setup can be ents only pressure

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].

ition all presented loss results are ass flow avthmust be determined to compute the coefficient. Hence, the static inlet pressure, the total inlet pressure and the total temperature in the settling chamber were measured. Furthermore the total pressure distribution was measured with two wake rakes 40% of chord length c behind the traicircumferential direction. One rake consists of 26 tside 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

reference coordinates system Figure 4: Cross section of the mounted cascade with

1q21 ),(

),(zupp

zu tt −=ζ

e cascade blades, which can

Figure 5: Total pressure loss parts

The total pressure loss copassage is computed 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 ζBL, the profile losses ζP caused by the friction of flow around the vane profile with infinite aspect ratio and the losses due to corner separation ζSP [1], [4], [8].

SPBLPges ζζζζ ++=

Experimental Errors

Systematic errors of the experiassessed. During the experimand temperature were measured. The pressure measurement chain consists of probes, transducer and voltmeter. The total pressure probe error

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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.

relative vane height z/h [-]

tota

lpre

ssur

elo

ssco

effic

ient

0 0.2 0.4 0.6 0.8 1

(upstream view)

(mass flow averared)

outfl

owan

gle

[°]

rela

tive

casc

ade

pitc

hu/

t[-]

0 0.2 0.4 0.6 0.8 1-0.5

0

0.5suction side

pressure side

lokal total pressure loss coefficient (mass flow averaged)ca

scad

epi

tch

u/t[

-]

0

0.5suction sidelokal total pressure loss coefficient (mass flow averaged)

Figure

trailing edge

relative vane height z/h [-]

tota

lpre

ssur

elo

ssco

effic

ient

0 0.2 0.4 0.6 0.8 1

(upstream view)

(mass flow averared)

outfl

owan

gle

[°]

rela

tive

0 0.2 0.4 0.6 0.8 1-0.5

pressure side

Figure

6: Measurement results at design point (Ma1= 0.66)

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7: Measurement results at incidenceof -6° (Ma1= 0.66)

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.

By means of fthe typical cothe suction sidetected. Thedue to incidencthe flow visuaseparation bubwhich give us the vane. Upslayer of the blturbulent.

In a second seat varying incand three inflo0.7) to obtain The peak efficthe curve whimeasuring andevices.

Figure Figure

8: Flow visualisation on the vane at design point (Ma1= 0.66)

low visualization (figures 8 and 9) rner separation on the rear part of de of the reference blades can be displacement of separation areas e from design point is detectable in lization figures too. Furthermore a ble on the suction side is shown information about the transition on tream of the bubble the boundary ade is laminar and downstream it is

ries measurements were carried out idence angles between -8° and +8° w Mach numbers (Ma1= 0.5 / 0.66 / loss curves as shown in figure 10. iency of the cascade is identified at ch presents a reference for further d development of flow control

Figure 10: Meafor v

ζ ge

s [-

]

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9: Flow visualisation on the vane at incidence of -6° (Ma1= 0.66)

surement results depending on inflow angle arious Mach numbers

inflow angle β 1 [°]

total pressure loss Ma = 0.7 total pressure loss Ma = 0.66total pressure loss Ma = 0.5

peak eff iciency- incidence +

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 784000 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].

incomingflow

leading edge trailing edge

laminar separation turbulent reattachment

area ofcorner

separation

Figure

Numerical Re

The incoming floware based on cascaBerlin. For the stacarried out at desiturbulence model Ghannam & Shawand the transition surface. In figurdistribution and stvane is shown whto peak suction. Tthe suction side avisible in figureconfirmed by the stransition coefficiepoint of laminarcoefficient cf beco

11: Computed static pressure distribution and strike lines on suction side of the

vane at design point (Ma1= 0.66)

sults

parameters for the simulation de measurements at the DLR in tionary computation, which was gn point of the cascade the k-ω and the Drela modified Abu

transition model was used [14] was allowed on the whole blade e 11 the computed pressure rike lines on suction side of the ereas the blue area is equivalent he laminar separation bubble on s indicated in figure 8 is also 11. This interpretation is kin friction coefficient cf and the nt shown in figure 12. At the separation the skin friction mes negative and the transition

Figure 12: Computed static pressure distribution P, skin friction coefficient cf and transition coefficient TF at mid span (at design point, Ma1= 0.66)

X

CF P

TF

0 .0 1 0 .0 2 0 .0 3

PC FT F

leading edge tra iling edge

lam inar separation turbulent reattachm ent

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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.

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 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 13: Comparison of computed and measured outflow angle distribution at design point (Ma1= 0.66)

0 0,1 0,2 0,3 0,4 0,5

relative vane height z/h [-]

outf

low

ang

le β

2 [°]

TRACE ComputationMeasurement

The experimentechanism of co

he wall and bladyer flow is dec

e backflowf separation relorner separationontinuing meas

Figure 14a

Figure 14b

Discussio

mtlafor thocc

: Computed local total pressure loss coefficient distribution atdesign point (Ma1= 0.66)

al results as well as the numerical results allow a grner separation. It is well known that this separation me boundary layer and the high pressure gradient in flow

elerated by the increasing pressure till separation. The n areas in both corners which cannot be obtained easily bative to the pressure distribution is known. The experi losses are an important part of the total losses of a cas

urements.

n of the Results

8

: Measured local total pressure loss coefficient distribution at design point (Ma1= 0.66)

ood assessment of the fluid mechanic ainly caused by the interaction between direction. The low energetic boundary

umerical simulation yields detailed data y experiments. Furthermore the position mental results have also shown that the cade and they will serve as reference for

Conclusions E on compr the

v ach number

In addition a n carried out. The comparison easured 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.

e and development mechanism of corner losses was achieved.

swertung von Nachlaufmessungen an ebenen Schaufelgittern, Bericht 67 A 49 AVA

[3] e

[4] he nd

Thermodynamics, Prague, March 2003

und Auswerteverfahren für ebene Verdichtergitter mit lussung bei hohen Unterschall- Geschwindigkeiten, DLR-IB-92517-02/B5, Dept.

[6] SCHEUGENPFLUG H.: Theoretische und experimentelle Untersuchungen zur Reduzierung der uste hochbelasteter Axialverdichter durch Grenzschichtbeeinflussung, PhD Thesis,

eswehr München, 1990; Germany

ittern, DK ,1954

r

[11] on des Schaufelseitenverhältnisses, PhD Thesis, Universität

[12] Cologne, February

xperimental investigations of secondary flowalues of losses at different incidences and M

umerical investigation was

essor cascade blades were performed. Therebys were determined.

between computed and m

Finally a better understanding of structur

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.

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