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Control of Compressor Face Total Pressure Distortion on a High Bypass Turbofan Intake using Air-Jet Vortex

Generators

S.D. Erbslöh* and Dr. W.J. Crowther† The University of Manchester, Manchester, UK, M13 9PL

J.R. Frutos‡ FEMTO-ST Institute, LPMO Department, 25044 Besançon, France

The control of turbofan intake duct flow separation and the subsequent control of compressor face distortion have the potential to achieve a relaxation of the maximum crosswind constraint at take off. Wind tunnel tests have been conducted using a 10% axis-symmetric scale model. Air jets, simulating micro-electrical mechanical actuators, have been implemented at the intake lip to reenergize the boundary layer locally. Improvements in distortion of up to 40% could be achieved at a 10% engine bleed.

Nomenclature at = average speed of sound at the axial position of minimum area , m/s Ah = duct area at leading edge, m² At = duct area at the axial position of minimum area, m² Aw = cross-sectional area of wind tunnel working section, m² A∞ = area of captured stream tube, m² Cp = local pressure coefficient at model surface = Pstatic/(1/2ρ(M tat)²) CR = contraction ratio = Ah/At d = radial distance from windward wall along the duct centerline, m D = diameter of the duct at compressor face, m Dh = diameter of the intake duct at the leading edge, m DC(60) = distortion parameter = (P f - Pθ=±30°)/Pf mjet = mass flow through air-jet vortex generators, kg/s mengine = mass flow through intake, kg/s Mj = average Mach number of an individual air jet at the orifice exit Mt = average Mach number at the axial position of minimum area Pa = ambient pressure, N/m² Pf = mean total pressure at the compressor face, N/m² Po = discrete averaged total pressure reading at the compressor face, N/m² Pstatic = discrete static pressure measured at the model surface, N/m² Psup = differential supply pressure for air-jet vortex generators , N/m² Pθ = mean total pressure at the plane of the compressor face included by the angle theta, N/m² Re = nacelle Reynolds number = ( ) µρ hDU ∞∞ s = surface distance from the leading edge along the internal duct, m smax = surface distance from leading edge to compressor face, m V∞ = average free stream speed, m/s

* Postgraduate Research Student , Goldstein Aerodynamics Research Laboratory / School of Engineering, Email: sascha.d.erbsloeh@stud.man.ac.uk, MAIAA † Lecturer, Department of Aerospace En gineering, Email: w.j.crowther@man.ac.uk, MAIAA ‡ Postgraduate Research Student, Institut FEMTO-ST, Laboratoire de Physique et Metrologie des Oscillateurs , jean-renaud.frutos@lpmo.edu

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Vt = average air speed at the throat, m/s β = angle of attack of intake model to the free stream, degrees µ = average viscosity of air, kgm-1s -1 ρ∞ = average free stream air density, kg/m³ θ = angle defined in the plane of the compressor face, degrees

I. Introduction he work presented in this paper is part of the AEROMEMS II research project sponsored by the European Union and a number of principal European aerospace contractors. It comprises of academic and industrial

partners specialized in a variety of disciplines including, micro machining and micro fabrication, experimental and computational fluid dynamics, turbo fan manufacture and flight vehicle systems integration. A key attribute of the present work is the close collaboration between micro manufacturing and systems engineering with wind tunnel investigation studies. Information flow in both directions ensures that the micro-electrical mechanical systems (MEMS) being developed are fit for purpose and also that experiments are based on technology that is achievable within the next year. This paper focuses on the application of MEMS flow control on a Pitot type intake system for a pod mounted high bypass turbofan engine typical for a transonic passenger airliner. The target flight condition is take -off with a high cross wind component. Presently, under these conditions, it is often necessary to limit the engine thrust such that major separation on the windward intake lip is avoided and the fan face total pressure distortion is kept within acceptable limits. Clearly, this has an adverse affect on the aircraft take-off performance, with available thrust at brake release being significantly reduced under high crosswind conditions. The intake design is hence compromised due to this off design point in the sense that the inlet lip has a larger leading edge nose radius than required at cruise condition. The primary goal of the present work is to develop an active system for controlling inlet separation and hence fan face total pressure distortion. The system needs to be active in the sense that it can be activated on demand on take-off, however, at all other times has minimal impact on overall intake performance. The implementation of a flow control system may benefit the inlet design either by relaxing the cross wind at take off constraint, i.e. allowing a higher cross wind with the same geometry, or to allow a smaller leading edge nose radius with the same critical cross-wind component.

So far several flow control methods have been considered for intake systems. The effect of inlet lip geometry1-8 has been studied thoroughly and several fluidic flow control methods have been reported. Many methods explored are based on mechanical alterations to the lip geometry, e.g. blow in doors as on the Harrier or variable chin inlet lip deflection as found on the Eurofighter Typhoon.9-11 Another class of methods is based on tangential blowing at the highlight or in the diffuser.12-15 There is also considerable work at present considering the use of flow control to manage the flow in ‘S’ ducts typically found on low observable applications and is normally focused on reducing swirl as well as reducing the total pressure distortion due to flow separation.9,10,16 Renewed interest has been shown to implement the S-Duct for an immersed inlet system on the blended wing configuration. 17,18 Though fluidic flow control methods have been proven successful in wind tunnel investigations none have found application on modern civil turbofan engines operating at present.

T

Figure 1: Sketch of diffuser and compressor cone

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Table 1: Test conditions

tM ∞V (m/s) tV

V∞ hA

A∞ ∞A

Aw

0.3 18 0.18 5.1 18.4 0.6 18 0.09 7.9 11.9

Figure 3: Variation of total pressure distortion with mass flow rate (graph by F. Aimar)

II. Aim The goal of the present investigation is to provide a flow control system that enables relaxation of the maximum

cross wind at take off constraint for civil turbofans. The system should be such that there is minimal cruise drag penalty due to the integration of the MEMS actuators . The flow control system should be reactive in that it can be turned on when required and ideally part of a closed loop system that actively detects separation and initiates an appropriate control response automatically.

III. Experimental Setup and Procedure This study considers an axis -symmetric intake of approximately 10% scale with a contraction ratio of CR=1.33 which corresponds to the side lip geometry of a full scale intake. The diameter of the duct at the station of the compressor face was 200mm. Figure 1 clarifies the definition of parameters chosen.

Tests were performed in the 2.8 by 2.1 meter subsonic wind tunnel of the Goldstein Laboratory at the University of Manchester. This closed return wind tunnel has a maximum speed of 70m/s and turbulence levels of around 0.5%. The inlet was tested at two different mass flow rates corresponding to M t=0.3, subsonic lip separation, and Mt=0.6, supersonic lip separation, with a constant free stream speed of 18m/s. Figure 3 illustrates these two operating conditions. The cross wind of 18m/s was found from industrial experience to be the maximum allowable free stream speed for the full scale intake-engine system at β=90°.

The model was connected to a reservoir and operated in suction mode resulting in an effective operating time of thirty seconds at Mt=0.3 and 15s at Mt=0.6. Two Venturi contractions were used to give the required mass flow rates. The model was mounted centrally inside the working section and the angle of attack (AOA) to the free stream could be adjusted from 0°<β<90°. A photograph of the model inside the working section is shown in Fig. 4 and in Fig. 5 a side view is sketched illustrating the individual components. Table 1 summarizes the fixed parameters at these two test conditions, which were carried out at a nacelle Reynolds number of approximately 0.4x106 in comparison to 6x106 for the full scale inlet.

The first flow control device to be implemented was the classical vane vortex generator (VG), used first to serve as a baseline and guide the placement of fluidic control devices. In the next step steady air jets were employed at positions identified from the VG tests. Air jets were arranged in a single row extending 220° circumferentially. The individual orifice was pitched 45° and skewed 90° to the local flow direction. These injection conditions have been optimized within the AEROMEMS II consortium, but similar conditions have been reported in the past.19-22

The orifices were machined to a diameter of 0.5mm the working material being aluminum. Each orifice was 2.5mm in thickness and hence had a relatively large thickness to diameter ratio of 5. This ratio should be kept to a minimum as the resulting vorticity of an air jet is dependant on the average velocity of the discharging jet.23 An undeveloped velocity profile is hence desirable.

A number of different actuator concepts may be used to achieve the required control input and the synthetic jet has attracted generous attention in recent years.24 Although actuator systems are usually relatively simple in their general layout, micro machin ing them is still posing a challenge and both aspect are being addressed in the Aeromems II project .25,26 In addition, special attention has to be given to micro fluid flow, which does not behave according to ideal flow theory.27,28 A simple example is the flow through small orifices. These experience large pressure losses if the diameter is smaller than a few hundred micrometers.

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Figure 4: Intake model inside the working section, free stream from right to left

Figure 7: Flow topology for separated duct flow at β=90°, Mt=0.6, V∞=18m/s; S1=primary separation line, F=focal point, S’=saddle point; free stream from top left to bottom right

Figure 5: Sketched assembly of intake model

In the AEROMEMS II project significant effort is underway to realize robust MEMS devices with an ongoing discussion between the demonstration projects and the manufacturing side to ensure realistic actuation and sensing equipment that will be able to provide the necessary velocity ratios or sensitivity. Fluidic vortex generators were chosen for the actuator type with a mechanism embedded giving a pulsating flow. Micro fluidic experiments have also been carried out to measure the jet velocities achievable through various orifice shapes and dimensions for varying pressure differences28 and supporting CFD studies and theoretical modeling have been completed. Special consideration was on developing electrostatic actuation as it offers good scaling properties to small dimensions, high energy densities and relative ease of fabrication as previous electrostatic actuators were manufactured during the AEROMEMS I project.29

Operating conditions and dimensions of orifices used in this project are expected to be achievable by MEMS actuators within the next year.

IV. Characterisation of basenline intake Surface oil flow visualization was used to

characterize the flow topology inside the diffuser. A mixture of kerosene and ultra-violet luminescent powder was applied to the surface evenly in the form of a fine spray. This method was also used by Hurd (1976)30 who observed a highly three dimensional flow field as sketched in Fig. 6. The surface topology in Fig. 7, obtained by the authors, consisted of a primary separation line at the leading edge (S1) , two focal points (F) placed symmetrical on either side of the duct and a secondary separation line at the top of the reversed flow region. A saddle point (S’) was positioned downstream of the focal points.

Figure 6: Sketch of surface shear lines for separated duct flow, face on view, after Hurd (1976)30

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Table 2: Distortion coefficients at β=90° and V∞=18m/s

tM DC(60) 0.3 0.022 laminar 0.6 0.058 laminar 0.6 0.023 turbulent

To investigate the performance of the baseline intake geometry an area weighted total pressure rake was installed at the same axial position as the compressor face and could be set at different circumferential angles to obtain a total pressure map of the compressor face. In addition, 19 static pressure taps were placed around the intake lip. Their position could also be changed circumferentially by rotating the diffuser. First the focus was on the centerline total pressure loss at different AOA for θ=0°. The duct flow was attached up to β=30° and fully separated at β=40°. Surface mounted transition devices were investigated for the test condition at Mt=0.6 and it was decided to place a roughness strip ahead of the primary separation line. Laminar separation at Mt=0.3 was assumed and the surface finish kept as clean as possible. The total pressure distortions are given in table 2 in terms of the parameter DC(60).

The static pressure distribution over the windward lip is shown in Fig. 9. For separated duct flow at β=90° the distribution has a distinct peak just after the leading edge, an abrupt increase in pressure and a flat distribution along the diffuser. For attached flow, in this case at β=30°, a distinct suction peak is also evident but is followed by a gradual pressure recovery along the diffuser. The static pressure shown in Fig.10 varied noticeably with circumferential position inside the diffuser, but the change in the local static pressure at the most upstream measurement was small. This was the position of the air jets and it hence may be assumed that these have a nearly constant discharge over their circumferential extent.

Figure 9: Static pressure distributions for separated (β=90°) and attached duct flow (β=30°) at Mt=0.6 and V∞=18m/s

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Figure 11: Lip separation at β=90°, Mt=0.6, V∞=18m/s; co-rotating VG array shown

Table 3: Improvements in DC(60) for a counter-rotating array, α=90°

tM DC(60) ∆ DC(60) Baseline 0.3 0.012 - 45% laminar 0.6 0.015 - 74% laminar 0.6 0.015 - 35% turbulent

Figure 10: Static pressure distribution at different circumferential positions; M t=0.6, β=90°, and V∞=18m/s

V. Distortion reduction using vane vortex generators A single array of vane vortex generators31-34,

triangular in plan form, was placed 15 vane heights upstream of the separation line and hence positioned just outside the leading edge over a circumferential extent of 180°. It should be noted that the position of the VGs was severely restricted by the surface curvature and the dividing stream line on the nacelle. The ratio of boundary layer thickness to vane height was not known but can be expected to lie in the range of 1-5. The vanes were hence not of the sub-boundary layer type but the smallest realizable dimension. Industrial CFD calculations have estimated the boundary layer at the leading edge of the full scale engine inlet to be about 2mm.

Figure 11 shows a surface flow pattern with a face-on view onto the leading edge. VGs are glued onto the nacelle with the direction of flow close to the surface being from bottom to top. The primary separation line is indicated by the dashed line. A co-rotating array is shown. Streaks show the trails of the individual vortices produced by the generators. At β=90° the separation line extended θ=±90° circumferentially.

Table 3 shows the effectiveness of the counter-rotating array in terms of improvements in DC(60) and Fig. 12 the contours of averaged total pressure at the compressor face. The region of high distortion at the windward wall is reduced noticeably in the circumferential and radial extent.

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VI. Distortion reduction using air jets Air-jet vortex generators were implemented at the same position as the vanes, 30 orifice diameters upstream of

separation. 82 jets, 0.5mm in diameter, were machined over a circumferential extent of 220° at a skew angle of 90°, a pitch angle of 45°, and a jet spacing of 5mm (10 jet diameters) resulting in a co-rotating array. The jets were operated with an internal plenum chamber incorporated into the intake lip geometry and could be supplied with a differential pressure of up to 1 Bar. Special consideration has been given to the machining of these orifices which will be used solely in this investigation to simulate embedded MEMS actuators. Different methods35 may be used to successfully machine orifice diameters larger than a few hundred micrometers, but using conventional drilling does not have any disadvantages with respect to surface finish. The achievable surface finish is of the order of 5µm, which should have space for improvement through optimization of the machining process.

The jet array gave a marked reduction in centerline total pressure loss for increased blowing pressure. Figure 13 emphasizes that the distortion is close to the windward wall as is also evident in Fig. 12 . The variations in centerline total pressure loss may be collapsed on a single graph by numerical integration. To obtain a non-dimensional form of the area Eq. (1) is used and referred to as a 2D distortion coefficient (DC2D). This parameter is convenient at this stage of the investigation as it allows the comparison of a large number of different test conditions with only a single measurement for each test case.

[ ]∫

−=

1

0

21

Dd

dPDP o

a

DC D (1)

The reduction in DC2D is shown in Fig. 14 for two different AOA and the trends displayed are typical for almost all test cases in the sense that 2D distortion decreases nearly linearly for increased blowing pressure. This linear behavior most probably indicates that a saturation of control effectiveness is not reached. This saturation seems to be the case under a reduced free stream of 9 m/s as DC2D shows to have an increasing gradient with higher blowing pressure. Figure 14 demonstrates that the reduction in distortion for Mt=0.3 is of the same order of magnitude. The jet spacing was also changed to 20 jet diameters (10 mm) for the test case at β=90° and suggests that the same control effectiveness may be achieved at half the mass flow rate, i.e. at twice the efficiency. Jet spacing and orifice diameter are the two geometrical parameters which may be changed with relative ease and within practical constraints. Figure 15 hence highlights how the efficiency of actuation may be raised through optimizing one of these parameters. In comparison to a counter-rotating VG array the jet array resulted in a control effectiveness of the same order of magnitude. A t low AOA the VGs were more effective and produced a lower distortion.

a) b) Figure 12: Averaged total pressure map at the compressor face; a) turbulent lip separation for unactuated model, DC(60)=0.023; b) single array of counter-rotating vanes, DC(60)=0.015; M t=0.6, α=90°, and V∞=18m/s; compressor hub not shown; free stream from left to right

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Figure 14: Improvement in 2D distor tion with increased differential blowing pressure for Mt=0.6 and V∞=18m/s

Figure 13: Change in Centreline Total Pressure loss with increasing supply pressure for M t=0.6, β=90°, and V∞=18m/s

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VII. Future Work The next phase of this project is to optimize the air jet array and achieve efficient actuation with the lowest

possible bleed air. Following detailed distortion measurements further efforts will look into the design of an appropriate control system and effective detection of flow separation. CFD calculations will be used to gain further insight into the flow physics around the leading edge and to be able to scale the physical dimensions of the VGs and jet array to the boundary layer properties.

VIII. Conclusion An air jet array in its initial design phase has been able to reduce total pressure distortion by as much as 40% at

the compressor face at 10% engine bleed. Its effectiveness was compared vane vortex generators and the improvement in distort ion was shown to be of the same order of magnitude. This suggests that relaxing the cross wind at take off constraint is feasible with an air jet array placed at the leading edge.

IX. Acknowledgments The AEROMEMS II project is a collaboration between BAE SYSTEMS, Dassault, Airbus Deutschland GmbH,

EADS-Military, Snecma, ONERA, DLR, LPMO, Manchester University, LML, Warwick University, TUB, Cranfield University, NTUA, and Auxitrol. The project is funded by the European Union and the project partners.

The authors would like to acknowledge the dedication and hard work of the technician team at the Goldstein laboratory of the University of Manchester.

Figure 15: Improvement in 2D distortion with increased differential blowing pressure for M t=0.3 and V∞=18m/s

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