Control of buffet phenomenon on a transonic swept wing

P. Molton, J. Dandois, A. Lepage, V. Brunet, R. Bur Control of buffet phenomenon on a transonic swept wing

AIAA Journal , Vol. 51, No. 4, pp.761-772, 2013

110

Control of Buffet Phenomenon on a Transonic Swept Wing

P. Molton,^∗J. Dandois,^†A. Lepage,^‡V. Brunet,^§and R. Bur^¶ ONERA–The French Aerospace Lab, 92320 Châtillon, France

DOI: 10.2514/1.J051000

The control of the transonic buffet phenomenon over a half-model with a swept wing has been investigated in the S3Ch wind tunnel of the ONERA Meudon center in the framework of ONERA’s joint research project BUFET’N Co.

A fine description of the buffet characteristics in the uncontrolled configuration is first given as a reference. Various steady and unsteady measurement techniques on the model surface and into the flowfield are used to perform a detailed analysis of the buffet phenomenon. The flow separation between the shock foot and the wing trailing edge being the cause of buffet, mechanical, continuous, and pulsed fluidic vortex generators are located upstream of the shock foot to reduce the extent of this separated area. As far as fluidic actuators are concerned several momentum coefficients, spanwise spacing, and orientations of the jets have been tested, while the effect of the actuation frequency is studied for the pulsed jets. Fluidic actuators, as well as mechanical ones, have been proved to be very efficient to postpone the buffet onset.

Nomenclature b = span of the model, mm Cp = wall pressure coefficient C_μ = momentum coefficient c = local airfoil chord length, mm

c_m = mean aerodynamic chord length, 220 mm d = hole diameter of the fluidic vortex generators, mm f = frequency of the pressure signal, Hz

h = height of the mechanical vortex generators, mm l = length of the mechanical vortex generators, mm MSVG = Mach number for the fluidic vortex generator M0 = freestream Mach number

pst = freestream stagnation pressure, Pa q = mass flow rate,g·s⁻¹

q0 = freestream dynamic pressure,ρ0U²₀∕2, Pa

Rec = Reynolds number based on the mean aerody-namic chord of the wing

rms-U = standard deviation of the longitudinal velocity component,m·s⁻¹

St = Strouhal number

T_st = freestream stagnation temperature, K U = longitudinal velocity component,m·s⁻¹ X,Y,Z = Cartesian coordinates, mm

α = angle of attack for the model or pitch angle for the fluidic vortex generators, deg

β = skew angle of the vortex generators, deg δ = boundary-layer thickness, mm

λ = spacing between two consecutive vortex generators, mm

I. Introduction

T

This phenomenon is a global flow instability known as“buffet”and can further lead to structural vibrations (“buffeting”). Buffet results in lift and drag variations that greatly affect the aircraft aerodynamics and, as such, it limits aircrafts’ flight envelope [1,2], because a marging of 30% with the lift coefficient of cruise conditions has to be respected by design standards. To postpone this buffet onset ONERA has launched in 2007 an internal research project named BUFET’N Co [3]; its final goal is to perform a closed-loop control of the buffet phenomenon.

A previous experimental study on a two-dimensional OAT15A profile [4] allowed a precise description of the aerodynamic condi-tions for buffet onset and a characterization of the periodic motion of the shock wave classically observed in two-dimensional buffet. The experimental results also served as a well-documented test case for the computational fluid dynamics community to validate advanced computing methods [5].

The main goal of the present study is the open-loop control of buffet on a half-model with a swept wing. This paper presents the description of the buffet phenomenon and the test of several control devices to delay its onset. Because the separation of the boundary layer is at the origin of buffet the objective of this experimental study is to postpone this buffet onset by suppressing or decreasing the separation zone by using control devices.

There is a lot of literature on the control of the shock/boundary-layer interaction. The control methods can be gathered into two main categories. In the first category, the objective is to weaken the shock by splitting it to have a bifurcatedλshock structure. Several studies in the last decade have examined passive control devices to bring about the modified shock pattern: a cavity covered with a perforated plate [6] and grooves and streamwise slots [7,8] underneath the shock foot.

These various concepts have led to a mitigation for success, reduction in wave drag being sometimes negated by viscous penalties [9]. This can be alleviated by using active devices, such as boundary-layer suction through a slot, but these devices required auxiliary equip-ment, which offsets any drag reduction benefits [10,11]. A promising method to lowering the total pressure loss through the shock system is the control by a bump. In the beginning, two-dimensional-shape bumps were investigated and led to significant wave drag reductions with moderate viscous penalties, but they were found to perform very badly at off design conditions [11,12]. More recent studies were performed with three-dimensional bumps, which have a limited Presented as Paper 2010-5494 at the 40th Fluid Dynamics Conference and

Exhibit, Chicago, IL, 28 June 2010–1 July 2010.; received 13 October 2010;

revision received 25 September 2012; accepted for publication 13 October 2012; published online 19 February 2013. Copyright © 2012 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Copies of this paper may be made for personal or internal use, on condition that the copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 1533-385X/13 and $10.00 in correspondence with the CCC.

*Research Engineer, Fundamental and Experimental Aerodynamics Department, 8 rue des Vertugadins, 92190 Meudon, France; Pascal.Molton@

onera.fr.

†Research Engineer, Applied Aerodynamics Department, 8 rue des Vertugadins, 92190 Meudon, France; Julien.Dandois@onera.fr.

‡Research Engineer, Aeroelasticity and Structural Dynamics Department, 29 Avenue de la Division Leclerc, 92322 Châtillon, France; Arnaud.Lepage@

onera.fr.

§Research Engineer, Applied Aerodynamics Department, 8 rue des Vertugadins, 92190 Meudon, France; Vincent.Brunet@onera.fr.

¶Research Engineer, Fundamental and Experimental Aerodynamics Department, 8 rue des Vertugadins, 92190 Meudon, France; Reynald.Bur@

onera.fr.

761

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spanwise extent, to enhance the off design performance [13–15].

Theλ shock structure has been found to propagate between the bumps, giving total pressure decreases across the span. Moreover, streamwise vortices developed along the bump sidewalls have a beneficial effect on the downstream boundary-layer behavior, rendering this passive control device as a promising concept.

The second category aims to energize the boundary layer upstream of the shock making it more resistant to the adverse pressure gradient and, consequently, less likely to separate downstream of the shock.

Mechanical vortex generators (VGs) [16–22], fluidic vortex genera-tors, and synthetic jet fall into this category. Previous studies done at ONERA [23] have shown that mechanical VGs are able to delay the buffet onset to higher angles of attack. However, even if they have demonstrated their efficiency for buffet onset delay, mechanical VGs have the drawback to increase drag in nominal cruise conditions.

This is the reason why fluidic VGs, which can be turned off, are also investigated. Moreover, they can be used in a closed-loop strategy to optimize the flow control, and consequently, to reduce auxiliary equipment for actuation, which is an important constraint for aircraft manufacturers.

Concerning fluidic VGs they have first been studied in [24] in the early 1950s. Most of the recent investigations have been aimed at optimizing the jet orientations. Rao [25] has performed an experi-mental study to control shock-induced boundary-layer separation with fluidic VGs. He has found that the optimum skew angleβis 45 deg, and the pitch angleαshould lie between 30 and 45 deg.

Pearcey et al. [26] have performed experiments of shock-induced separation control by fluidic VGs for the freestream Mach number between 1.25 and 1.65. They observed that only one streamwise vortex forms for a velocity ratio value, which depends on the skew angle and is minimum forβ45 deg. The optimum angles found wereα45 degandβ60 deg. Szwaba et al. [27] have realized a genetic algorithm optimization of pitch and skew angles at a Mach number of 1.25. They found that the vorticity is maximum for α20 degand β70 deg. Seifert and Pack [28] have used a synthetic jet slot upstream of the shock on a NACA-0015 at M0.55. They have observed in the wake a reduction of the pressure fluctuation level associated with the buffet. Hassan et al. [29]

have performed two-dimensional unsteady simulations of a synthetic jet slot upstream of the shock of a NACA-64A010 airfoil at M0.85. They have observed a decrease of the total drag, which is due to the replacement of the single strong normal shock by a two-shock system, a weak oblique two-shock, and a weak normal two-shock.

Pulsed fluidic VGs have been studied in [30] on a two-dimensional DERA M2303 airfoil. The pulsed blowing system consists in a rotating shaft fixed in a pillow block. The pitch and skew angles were, respectively, equal to 45 and 90 deg. The Mach number of the jets was 2.94. In the continuously blowing case they have observed a lift increase by eliminating the shock-induced separation experienced at high-lift conditions but also a total drag increase lowering the lift over drag ratio. Drag decreases when the frequency of the pulsed jets increases.

In this paper, the studied configuration is a well-equipped half-wing/body configuration designed during the internal ONERA project BUFET’N Co. First, the experimental setup including the tested control devices and the measurement techniques is presented.

The results obtained for the uncontrolled configuration are detailed, especially flow characteristics at the buffet onset and for the well-established buffet regime. At last, controlled configurations, includ-ing mechanical VGs and continuous and pulsed fluidic VGs are evaluated and compared to the clean configuration in order to draw some conclusions on the efficiency of these actuators to control the buffet phenomenon.

II. Experimental Setup with Control Devices and Measurement Techniques

The study is carried out in the continuous closed-circuit transonic S3Ch wind tunnel of the ONERA Meudon center. This facility is powered by a 3500 kW two-stage motor-ventilator group, equipped with 24 blades, which rotate at a fixed rotational speed of 1500 rpm.

The test section size is0.76 m×0.82 m×2.2 m. The Mach number domain extends from 0.3 to 1.2. The stagnation pressure is the atmospheric one, and the stagnation temperature lies between 290 and 310 K. The shapes of the upper and lower walls are adapted for each flow condition based on a steady flow hypothesis so as to reproduce far-field conditions. Side walls are equipped with schlieren quality windows. The flow velocity is adjusted by a downstream sonic throat, which warrants a 10⁻⁴ uncertainty of the Mach number value. The freestream Mach numberM0is set at 0.82 for all tested configurations. The angle of attackαof the model can be varied between 2 and 4 deg by a mechanical system and/or a proper adjustment of the adaptive walls. By continuously varyingαone could accurately find the value for buffet onset.

The experimental arrangement is shown in Fig. 1. The model is composed of a swept wing attached on a half-fuselage. This model was designed during the BUFET’N Co project, and most of the wing is based on the supercritical OAT15A airfoil. The swept angle at the leading edge is equal to 30 deg. The wing twist was adapted to ensure a constant pressure along the span in cruise conditions, as well as a shock parallel to the leading edge. From root to tip, the chord varies between 240 and 200 mm over a span of 704 mm. In the end, no separation at the wing root was ensured thanks to adapted profiles and twist in that region.

The Reynolds number based on the mean aerodynamic chord (c_m220 mm) isRe_c2.5×10⁶. Boundary-layer transition is triggered on the model by using a carborundum strip located at X∕c0.07on both the upper and lower side of the model, as well as on the fuselage. The model is equipped with 49 static pressure taps, 39 unsteady Kulite™pressure transducers, and 6 accelerometers.

Figure 2 shows their locations on the suction side of the wing.

Fig. 1 Experimental setup in the S3Ch wind tunnel.

Y/b = 0.5

Fig. 2 Locations of pressure taps, Kulite transducers, and accelerom-eters on the model.

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In this study, four control devices are investigated and four dedicated covers have been manufactured. The first control device is of passive control type, i.e., mechanical VGs (VGms). Because of the swept wing, only corotating VGs are considered here. The VGs, located at 20% of the chord, consist in 27 small triangles with a height hδ1.3 mmand a length equal to 5 h. Their skew angle has been defined thanks to numerical simulations [31] and is equal toβ 30 degwith respect to the freestream direction (and soβ0 deg with respect to the leading edge). The first VG is located 51% of the span (b), the last one at 89%, and the spacing between the VGs is 1.7% of the span (λ12 h). The other three control devices are active: two continuous fluidic VGs and one pulsed fluidic VG. The fluidic VGs consist in small nozzles with a supersonic exit flow at M_SVG2. The exit diameter of the nozzles (d) is equal to 1 mm, and the pitch angle (defined between the jet direction and the local wall tangent, see Fig. 3) isα30 deg. The 40 continuous fluidic VGs are located between 53% and 82% of the span with a spacing equal to 0.85% of the span (λ6 mm). The 25 pulsed actuators are located between 50% and 84% of the span with a spacing equal to 1.63% of the span (λ11.5 mm). The orientation of the jets with respect to the leading edge of the modelβis an important parameter; this is the reason why it has been studied numerically [31], in order to define the most interesting skew angles to be tested. Thus, two skew angles for continuous fluidic VGs have been tested (and are named VGF4 and VGF5) and one for the pulsed fluidic VGs (VGFp). These pulsed fluidic VGs consist in ONERA homemade piezoelectric actuators supplied with compressed air and driven by an electric square signal.

They are located at 23% of the chord. The characteristics of each VGs’configuration are summarized in Table 1.

For the fluidic VGs, the momentum coefficientC_μis defined by

C_μ ρjS_jU²_j

1

2ρ0SU²₀ qmUj 1

2ρ0SU²₀ (1)

whereρjandU_jare, respectively, the density and velocity of the jets;

S_jthe sum of all orifice surface area based on the hole diameter (not the projected surface); andqmis the mass flow rate. When the flow at the exit of the nozzles is supersonic the Mach number (M2) and, thus,Ujare fixed, and only the mass flow rate continues to increase with the air supply stagnation pressure. In the case of pulsed blowing the time-averaged mass flow rate and jet velocity are used in Eq. (1).

The variablesρ0andU0are, respectively, the freestream density and velocity of the main flow; the wing surface corresponding to a half-span being denotedS.

The first part of the study consists in researching the conditions for buffet onset by varying the incidence of the model and then in qualifying the reference flow in the presence of buffet. During this first stage, pressure measurements around the model were performed (steady and unsteady pressure), as well as visualizations of the flow by viscous coating. In the second part of the study, the effectiveness of various control types was evaluated using the same measurements.

The flowfields for the baseline and for some controlled configurations were measured using two components particle image velocimetry (PIV 2C) and three-components Laser Doppler Anemometry (LDV 3C) systems. The PIV system is composed of a pulsed2×400 mJNd/YAG laser with a wavelengthλ532 nm.

The acquisition of image pairs is performed by two PCO cameras (1376×1040 pixels). Each camera is equipped with an autofocus controller, a 50 mm Nikon lens, and an interference filter with a wavelengthλ532 nm. The management of the image acquisition is performed by the Davis software v. 7.2. The laser pulse rate is set at 4 Hz, and the time between two images at1.6μs. The acquisition time for one camera is about 260 s. The image processing is done using the PIV software FOLKI [32,33]. This software, optimized for massively parallel architectures like graphics processing units, produces dense vector fields with a very limited computing cost.

Moreover, for each configuration, a mask is used in order to not calculate velocity vectors in the model and hiding areas dazzled by the reflections of the laser plane. The flow is seeded with oil droplets emitted by a device implanted at the inlet of the diffuser downstream of the test section. The average size of the droplets is of the order of1μm. The measurements plan of each camera is presented in Fig. 4. Indeed, the first camera focuses on the development of the upstream flow until two-thirds of the chord, and the second camera focuses on the last one-third and the wake. The model deformation and the dazzled areas are represented by a black mask in Fig. 4, and the profile geometry in the measurement plane is represented by a white line. The dazzled areas are mainly due to the reflection of the laser plane on the model accentuated by the thin film of particles used for the flow seeding. The spatial resolution of the PIV fields is 0.11 mm both in thexandzdirections. The overlapping of the measurements plans ensures homogeneity of the results. Flow statistics are computed on 1000 instantaneous snapshots. PIV 2C measurements were carried out atY∕b0.6, corresponding to the section equipped with Kulite sensors, where the massive separation downstream of the shock is observable.

For the LDV 3C measurements, the optical setup is used in

“forward scattering”mode. On the emission side, the system consists Fig. 3 Sketch defining the VGs’parameters.

Table 1 Parameters of the different VGs’configurations

Configuration Control type Dimensions (mm) λ βdeg∕LE Y∕b(%) No. of VGs

1 Mechanical (VGm) h1.3,l5 h 12 h 0 from 51 to 89 27

2 Fluidic (VGF4) d1 6d 60 from 53 to 82 40

3 Fluidic (VGF5) d1 6d 30 from 53 to 82 40

4 Fluidic pulsed (VGFp) d1 11.5d 60 from 50 to 84 25

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of a laser Spectra Physics 2060, a fiber beam splitter, and two optical systems creating a volume of150μm. The optical receiver is based on Cassegrain telescopes collecting the light flux from each particle.

The acquisition of measurement signals is carried out by means of a TSI-IFA 750 system. The coincidence window is set to1μs. Raw recorded measurements are analyzed using the software ASSA developed at ONERA [34]. The flow is seeded by means of a generator located in the collector of the wind tunnel. The average size of the droplets is of the order of1μm. For this study, the mean and fluctuating velocity components are calculated by averaging over 5000 samples for each measurement point. Surveys were made vertically on the upper side of the model for different longitudinal positions. Each line consists in about 55 measurement points. The first point is at about 1 mm of the model. Measurement points are concentrated near the wall and in the separated zone and are more spaced in the external flowfield. Figure 5 shows the meshes used for each plan.

III. Results

First, the baseline (uncontrolled) configuration is analyzed to determine the flow conditions for which buffet occurs. The flowfield

First, the baseline (uncontrolled) configuration is analyzed to determine the flow conditions for which buffet occurs. The flowfield