American Institute of Aeronautics and Astronautics 1
Generic Transport Aft-Body Drag Reduction using Active Flow Control
Eli Ben-Hamou*
Faculty of Engineering, Tel-Aviv University, Tel-Aviv, ISRAEL
Eran Arad Aeronautical Systems, RAFAEL Ltd., Haifa, ISRAEL
Avi Seifert+
Faculty of Engineering, Tel-Aviv University, Tel-Aviv, ISRAEL
Active flow separation control is an effective and efficient mean for drag reduction and unsteady load alleviation resulting from locally or massively separated flow. Such a situation occurs in configurations where the aerodynamic performance is of secondary importance to functionality. The performance of heavy transport helicopters and planes, having a large, and almost flat, aft loading ramp suffer from the poor aerodynamics of the aft body. Hence, a combined experimental and numerical investigation was undertaken on a generic transport plane/helicopter configuration. The experimental study was composed of surface pressures and direct drag measurements and surface as well as smoke flow visualization. The baseline flow was numerically analyzed, using finite volume solutions of the RANS equations, in steady and time-accurate modes. The baseline flow around the model was insensitive to the Reynolds number in the range it was tested. The flow separating from the aft body was characterized by two main sources of drag and unsteadiness. The first is a separation bubble residing at the lower ramp corner and the second is a pair of vortex systems developing and separating from the sides of the ramp. Apparently, a secondary bubble on the ramp causes increased suction and elevated drag as the model incidence is being reduced from positive to negative angles. As the incidence decreases the pair of vortex systems also penetrates deeper towards the centerline of the ramp. As expected, the ramp lower corner bubble was very receptive to periodic excitation introduced from four addressable Piezo-fluidic actuators situated at the ramp lower corner. Total drag was reduced by 3-11%, depending on the model incidence. There are indications that the flow in the wake of the model is also significantly steadier when the bubble at the lower ramp corner is eliminated. The vortex system is tighter and steadier when the bubble is eliminated.
American Institute of Aeronautics and Astronautics 2
Introduction
Helicopter fuselage is rarely designed for low drag and high aerodynamic efficiency. Rather, the design process is usually dictated by functionality and operational aspects, not necessarily low drag. One such functional parameter is the aft loading ramp in large transport helicopters and planes. The presence of the aft door (or “Ramp” as it often referred to) dictates flat and inclined surface connected to the fuselage at almost its widest cross section. This, in turn, causes flow separation at the lower corner, generating a bubble, and it generates a system of stream-wise vortices at the sides. These vortical structures bear some resemblance to the delta wing post-stall vortices or to those separated from hatch-back cars. The recent success in controlling the flow over delta wings (Ref.1) as well as bluff bodies (Ref. 2) leads to the current study, aimed at reducing drag and alleviating unsteady loads resulting from separated flow at the aft body of a generic helicopter/transport plane configuration.
Experimental set-up
The experiments were conducted on a generic transport helicopter/plane body, similar to those shown in Fig. 1, with “details” of secondary importance to the current study removed The 250 mm by 250 mm cross section body (Fig. 2a) was cut at an angle of 30 deg to the lower surface, as in the C-130 transport plane and in the Ch-53 heavy helicopter (Fig. 1). The length (from the two extreme streamwise points) of the model is 1000mm. The model was installed in the Meadow-Knapp low-speed wind tunnel test section, at the Meadow aerodynamics laboratory of Tel-Aviv University, as shown in Fig. 2. The tunnel test section height is 1500 mm, its width is 609mm and the tunnel speed range is 4-65m/s. Turbulence level is below 0.2%. Twenty four individually operated compact, cavity installed, Piezoelectric fluidic actuators were positioned around three sides of the plate simulating the aft ramp surface (Fig. 3, “ramp” hereafter) and ejected the periodic excitation at about 45 deg in the downstream direction from 1mm (wide) slots. Note that actuator #1 is at the upper right corner while actuator #24 is at the upper left corner. The actuators were operated by computer controlled eight-channel arbitrary function generation system. The signals were amplified in a 50:1 voltage ratio (amplitude to rms). Feedback signals from all actuators were read by the tunnel data acquisition system for health monitoring and performance validation. The actuators were calibrated, as installed in the model, on a dedicated bench-top calibration rig. The Helmholtz resonance frequency of the 24 actuators was between 1.5 and 1.8kHz, and those are the working frequencies. Fig. 4 shows the amplitude response of four of the 24 actuators at the working frequency. The scatter in the peak velocity is roughly ±10% providing span-uniform excitation momentum coefficient within ±20% when using the same voltage input signal. The definition of the excitation momentum coefficient is based on the total excitation
momentum normalized by a reference free-stream momentum in the form: 2..
_
2∞∑= UAuAC SC
slotsactivepeakslotµ ,
where slotA is the excitation slot exit area and ..SCA is the model cross section area (=0.25 by 0.25 m2). The
dimensionless frequency is defined as: ∞+ = UfdF / , where f is the excitation frequency (1770Hz for all the
results presented here), the reference scale is the model width, d=250mm, and ∞U is the free-stream velocity (8 to 25 m/s in the current study). The resulting model length Reynolds numbers (Re) are in the range 0.5 to 2.0x106. A strip of double back tape, with roughness elements (rms diameter 1mm) is scattered on it to induce transition. No measurable difference was recorded either with total drag or Cp considered, indicating turbulent separation with small contribution to the total drag due to skin friction.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 3
Fig. 1 The similar aft bodies of the C-130 and Ch-53.
Fig. 2a A view of the test model as installed in the Meadow-Knapp wind tunnel. Note the strut entering the model from above. Curved surfaces close to the aft loading ramp shown in Fig. 1, were omitted for simplicity and generality.
Amplitude Scan - 1770 Hz - bottom set
0
5
10
15
20
25
30
0 10 20 30 40 50 60 70 80 90
Amplitude [VAC]
Velo
city
[m/s
]
piezo 11
piezo 12
piezo 13
piezo 14
Fig. 4 Amplitude scan of the four actuators installed at the lower Ramp edge, showing peak slot exit velocity at 1770Hz.
Fig. 3 A back view of the model showing the ramp and actuator slots. Actuator #1 is on upper right corner. There are 9 identical actuators on each vertical side, 2 corner and four lower side actuators.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 4
Figure 2b: Computational domain that matches the wind tunnel configuration.
The experimental model (Fig. 2a) was equipped with one row of streamwise pressure taps along its symmetry plane, in mid-width location, one row of pressure taps around the mid-height of the model and three rows of spanwise pressure taps located on the back cover plate (“ramp”). Pressures were measured by PSI Inc. pressure scanner with a resolution of 0.02psi. The small number of the pressure taps allows only very crude estimation of the forces and moments. However, conditions in which the baseline flow was separated could be identified and distinguished from cases where excitation had a significant effect on the flow. Drag force was directly measured by a load cell installed in the model.
Computational Model The physical phenomenon under consideration in essentially unsteady, with oscillations that are not restricted to one typical frequency. The flow behind the ramp, which is the region of interest from flow- separation control point-of view, is turbulent with several separation bubbles. This nature of the flow-field directs towards the use of time-accurate computational method. As the flow control apparatus intervenes with the mini-structures of the flow field and their detailed temporal development, direct numerical simulation (DNS) would be the most appropriate analysis tool. However- the substantial computational resources that are required by this approach make it less attractive for practical cases (with large Reynolds numbers) and complex configurations. The computational problem can be eased by the use of a computational model which simulates only part of the flow structure, and uses simplified models to account for the rest. Large-eddy-simulation (LES) and unsteady-Reynolds-Averraged-Navier-Stokes (URANS) solutions fall into this designation. In LES methods, only the small-scale flow-structure is carved out of the simulation, and is represented by a model (which can be of different levels of accuracy). The large-scale turbulent flow-structures, as well as the mean-flow mechanism, are directly simulated. URANS approach avoids the simulation of the whole unsteady-turbulent structures, and account only for the mean-flow structures. The whole contribution of the turbulent flow-structures is represented by a model. The time-accurate approach is added artificially to the already time-averaged equations (RANS). The difference in the level of modeling affects the amount of computational resources that are required by these two approximations in two-ways: First, proper simulation of large eddies requires a finer computational mesh than needed to account only for the mean-flow (RANS). This requirement is especially severe near solid walls, were
Inflow
Tunnel walls
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 5
the typical scales become very small. Furthermore, while URANS capture a single (or few) frequencies (e.g. coherent shedding), LES resolves the (nearly) complete range of scales. Consequently LES requires much longer integration time to build a statistically averaged solution. In the present study a combined approach is taken. The results which are presented in this paper were obtained using the solution of RANS and URANS equations. In the next phase, LES simulation, which has already been applied for flow-separation analysis (Ref 3), will be used to provide better representation of the temporal flow mechanism. The URANS equations were solved mainly by FLUENT commercial code, using an incompressible formulation. Both realizable k- ε and Menter k-ω SST turbulence model were used. A segregated solver was applied, using second-order pressure interpolation. SIMPLEC was used for pressure-velocity coupling. The Quick scheme was used for the turbulence models equations. Second order point-implicit scheme was used for temporal integrations. The LES code is an in-house program, and is described in Refs. 3 and 5. The solution domain, presented in Fig. 2b, was set to be identical to the wind-tunnel setup, discussed in the previous section. Hexahedral multi-block mesh was generated using ICEM-CFD mesh generator. Fine resolution behind the body, and wall resolution of Y+=O(1), necessitated a mesh of about a million cells. The computational facility was a cluster of X86 processors, with Myrinet and Gigabit switches. MPI library was used for parallel computation.
Discussion of Results
Baseline (uncontrolled) Flow Before attempting to control such a complex flow field, it is essential to understand the baseline. This was achieved using a combined experimental-numerical approach. The experiments included steady and unsteady wall pressures measurements at selected locations, as well as surface and smoke flow visualization. The 3D numerical simulations were performed with and without the tunnel walls present using RANS and URANS solutions (discussed above). The experiments revealed the presence of a separation bubble situated at the juncture between the lower side of the fuselage and the ramp. Tuft flow visualization showing separated flow region close to the lower corner of the ramp-aft body juncture is presented in Fig. 5. The separation bubble is discernible in the computed ramp skin friction, presented in Fig. 6. Here blue color indicates separated flow. It can be seen that computed flow structure is in good agreement with the tufts. The separation bubble extends about two thirds of the ramp width, at the juncture with the main body lower surface, and has an approximate triangular shape, closing in about mid ramp height. Figure 7 presents the streamwise pressure distribution on the upper and lower sides of the main body of the model. These experiments were performed at a negative incidence of –2.5°. The computational pressures coefficients on top of the model are slightly larger than the measured values, indicating a smaller acceleration. The pressure distribution on the ramp centerline is presented in Fig. 8. The results of a URANS (unsteady) calculation were averaged over a large time period. The full line represents the average, while the dash-lines designate the region of oscillations of the solution. The agreement between measurements and computed results in this comparison is reasonable, taking into account the fact that two-equation turbulence models do not excel in separated flow simulation. A more suitable model for separated flow, v2f, will be used in the next phase of the investigation. However, both approaches indicate that there are most probably two bubbles on the ramp: one immediately at the beginning upstream region of the ramp, and a second towards its trailing edge. The spanwise structure of these bubbles was examined using additional methods, and is discussed below. Regardless of the spanwise structure, it is expected from previous experience (e.g., Ref 4), that both bubbles will be most receptive to periodic excitation. The spanwise extend of the bubbles was determined from tufts flow visualization (Fig. 5) and calculated skin friction (Fig. 6) on the ramp. The lower bubble is easily identifiable in both figures. It extends about two thirds of the ramp width, at the juncture with the main body lower surface, and has an approximate triangular shape, closing in about mid ramp height.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 6
The secondary bubble was not seen in the tufts flow visualization of Fig. 5. The calculated skin friction (Fig. 6) indicates that it is situated close to the upper edge of the ramp and adjacent to its sides. As will be shown later, it is most probably connected to the two vortices separating from the sides of the ramp. Additional, perhaps dominant, flow features are the two vortex systems that originate at the lower corners of the ramp. These vortices were both measured and calculated. Surface oil flow visualization (Fig. 9) clearly shows the signature of the two vortices. The vertical sides of the ramp show separated flow regions, or wedges, with increasing width as one moves away from the lower side corners towards the upper parts of the ramp. These are surface signatures under the region of influence of the side vortices. Two regions where the oil flow was completely washed away, near the lower corners, indicate the reattachment region of the vortices. The region in between, where the oil flow was not affected, is the lower separation bubble. Figure 10 shows computed surface “oil flow” pattern. Placing lines of “dots” and allowing the skin friction to determine the direction in which the “oil” will migrate generated this pattern. The signature of the vortices is clear and is in agreement with the experimentally determined oil flow visualization shown in Fig. 9. Additional features of the side vortices will be discussed towards the end of the paper.
An additional insight into the structure of the lower separation bubble can be gained from the smoke flow visualization image shown in Fig. 11 and the computational pressure contours shown in Fig. 12. The combined consistent image shows that the bubble extends half the ramp length on the centerline. Based on our previous experience (e.g. Ref 4), this situation is extremely receptive to nominally 2D periodic excitation that is introduced at the separation region. The spanwise structure of the flow imprint on the ramp can be appreciated by considering the computed and measured ramp pressures, presented in Fig. 13. The time averaged Cp measured by the three spanwise rows of pressure taps in qualitative agreement with the computed pressures. The flow is nominally 2D in the central part of the ramp. The quality of agreement is poorer at the lower ramp region since the computation underestimates the severity of the separation. This tendency of the CFD model using a two-equation turbulence model was already discussed. The prediction is expected to improve by using a more suitable model, like v2f. The imprint of the side vortices could be identified as higher acceleration on the sides of the ramp.
Figure 5: Tuft flow visualization on the ramp. Figure 6: Computed skin friction on the ramp.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 7
0 0.1 0.2 0.3 0.4 0.5−1
−0.5
0
0.5
1
1.5
X [m]
−C
p
Compexp
0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45 0.5
−0.5
−0.25
0
0.25
0.5
0.75
1
1.25
1.5
1.75
X [m]
−C
p
ComputedExp.
Figure 7: Experimental and computational pressure distribution on top of the model centerline.
Figure 8: Experimental and Computational pressures distributions on the "ramp" centerline. Note that here x is measured from the Ramp corner, along the ramp.
Figure 13: Computed and measured Ramp pressures, U=15.8m/s.
Controlled flow To control the separation bubble located at the lower juncture between the aft body and the ramp, four Piezo fluidic actuators, situated at the lower ramp corner were used, each actuator slot was 40mm long and 1 mm wide. The actuators’ slots were situated on the ramp, about 1 mm from the corner, due to fabrication considerations. This caused the excitation be less than ideally situated as one would optimally desire, i.e., directly at the corner where the flow separates. The excitation emanated at an angle of about 45 deg to the ramp surface, or about 15 deg to the lower model surface and was directed in the downstream direction. The amplitude calibration of these four actuators, at the operating frequency of 1770Hz, was presented in Fig. 4. A typical peak slot exit velocity (measured at the slot exits using a hot wire) to excitation voltage relationship could be seen. Peak velocities exceed 25 m/s and the scatter between actuators is about ±7%. The baseline and controlled flow response at Re=1.2x106, α≈β≈0, is shown in Figs. 14a-14c. The baseline, uncontrolled bubble can clearly be seen at x=550-650mm using streamwise pressure distributions (Fig. 14a). When the excitation was activated, the lower ramp corner separation bubble was completely eliminated. The resulting effect was a reduction of about 7-10% drag and significantly steadier wake. The activation of the lower corner actuators had no measurable effect on the 2nd bubble, shown in both baseline and controlled flows at x=700-900mm (Fig. 14a). The spanwise, baseline and controlled, Cp’s on the ramp are shown in Fig. 14b, for identical conditions to those of the data presented in Fig. 14a. It can be seen that for both flow conditions, the pressure is almost span-uniform for the 150-175mm around the centerline. The excitation reduces the pressure close to the corner edges (possibly indicating tighter vortices), and increases it elsewhere. This is an indication of eliminating the bubble at the lower ramp corner, as shown by the side view of Fig. 14a. Figure 14c presents the time-resolved drag coefficient, measured directly by the balance installed in the model, at similar conditions to the above, with the addition of the two corner actuators (which later were found to have negligible effect at β=0). The clear reduction of about 7% in drag can be seen as the actuators were turned on. Perhaps not less important is the significant reduction (about 50%) of the drag fluctuations when the actuators operate. It needs be emphasized that these data were sampled at 0.1 sec integration time and that the resonance frequency of the load cell with model installed was about 7Hz. Additional data needs to be acquired to validate the above result and identify the physical mechanism, but the data trend warrants this further study.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 9
Cp - Cover P tap lines - 18 m/sec
-3
-2.5
-2
-1.5
-1
-0.5
0
-150 -100 -50 0 50 100 150z [mm]
Cp
High-Baseline
Middle-No Excitation
Low-No Excitation
High-bottom+corner
Middle-bottom+corner
Low-bottom+corner
Cp - vertical cut - 18 m/sec
-1.5
-1
-0.5
0
0.5
1
0 200 400 600 800 1000
X [mm]
Cp
Baseline
Bottom Actuors Operating
Fig. 14b Baseline and controlled Ramp pressures, measured while the four actuators installed at the lower Ramp edge operated at 1770Hz
Fig. 14a Pressure distributions measured around the model half width. The baseline shows two separated regions on the ramp, the lower was eliminated by the excitation emanating from the four actuators installed at the lower Ramp edge, 1770Hz.
0.47
0.48
0.49
0.5
0.51
0.52
0 30 60 90 120 150 180
Time [sec]
CdBottom set+corner
on
Fig. 14c Drag measured by balance, Showing Cd reduction by 7% and Cd Excursion reduced by 50%
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 10
Effect of Model incidence on Drag - baseline ; 20 m/s
Fig. 15b Baseline ramp pressures vs α, U=20m/s, β=0. Fig. 15c Baseline pressures along the ramp longitudinal centerline, U=20m/s
Additional tests conducted at several incidence angles (α) indicated the great sensitivity of the flow field and drag coefficient to variations in the incidence angle. Fig. 15a presents the total drag measured by the internal balance at zero slip angle (β). The same trend was noted at practically all tested slip angles (7.5deg >β>-7.5deg). A possible explanation to the increased drag as the incidence angle is decreased (i.e. nose down motion) can be extracted by analyzing the data presented in Figs. 15b-c. In Fig. 13nb, one can note that the signatures of the side vortices penetrate deeper towards the ramp centerline and the pressures become increasingly negative. This is an indication of stronger side vortices that will contain more angular momentum, implying higher drag. Also the more negative surface pressures indicate larger pressure drag. Fig. 15c presents the pressures along the ramp longitudinal canter-line showing increased suction due to what appears to be a larger secondary bubble on the upper half of the ramp. This trend also contributes to the growth in drag as the incidence decreases. Figures 16-18 present ramp pressures, flow visualization images and drag coefficients when the four lower ramp corner actuators control the lower ramp corner separation bubble for a range of incidence angles. At all incidence angles the excitation was effective in eliminating the bubble. Figures 16a and 16b show ramp pressures for the baseline and controlled flows. At both incidence angles considered, the elimination of the bubble is evident. Figures 17a and 17b show smoke flow visualization images for the baseline and controlled flows, respectively, focusing on the left side vortex system. The baseline vortex is quite diffused and unsteady. Periodic excitation introduced at the lower ramp corner significantly tightens and stabilizes the vortex (Fig. 17b), in addition to eliminating the lower bubble.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 11
Ramp Vertical P Taps - alpha=-0.5 deg ; 20 m/s
-1.3
-1.1
-0.9
-0.7
-0.5
-0.3
-0.1
0.1
0.55 0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1X/C
Cp
Cmu=0
Cmu=0.001
Fig. 16a Baseline and controlled ramp pressures, U=20m/s, α=-0.5 deg, control via 4 lower ramp corner actuators, total Cµ=0.4%.
Fig. 17a Baseline flow, only left corner vortex system visualized, U=20m/s, α=0
Ramp Vertical P Taps - alpha=1.2 deg ; 20 m/s
-1.3
-1.1
-0.9
-0.7
-0.5
-0.3
-0.1
0.1
0.55 0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1X/C
Cp
Cmu=0
Cmu=0.001
Fig. 16b Baseline and controlled ramp pressures, U=20m/s, alfa=1.2 deg, control via 4 lower ramp corner actuators, total Cµ=0.4%.
Fig. 17b Baseline flow, only left corner vortex system visualized, 4 lower actuators, U=20m/s, α=0
Fig. 18a Drag coefficient vs Cµ when control applied via 4 lower corner actuators
Fig. 18b Drag reduction (% of baseline drag) vs α, U=20m/s, lower corner actuations.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 12
The drag coefficient for two of the incidence angles tested is presented in Figs. 18. Note that the baseline drag is for Cµ=0. At both near-zero incidence angles, the drag reduction saturates at a total Cµ≈0.2% (the data shown in Fig. 18 is for Cµ of one actuator, i.e., one fourth of the total Cµ). The lack of additional effect at larger Cµ is because the bubble has been completely eliminated by this level of excitation and further increase is exciting already attached flow, i.e. wasted energy in the current context. Fig. 18b presents relative drag reduction, referenced to the baseline drag of Fig. 15a. It can be appreciated that drag reduction of 3-11% was measured, depending mainly on the level of baseline drag. All incidence angles considered showed drag reduction.
Figures 19 show baseline and controlled surface oil flow visualization corresponding to the data of Figs. 14 and 17. The main effect of the excitation is increased skin friction at the lower ramp corner and its side edges, as indicated by the “washed out” regions where the pigment has disappeared. An additional effect could be seen by a stronger and narrower imprint of the side vortices that also persists almost to the upper edge of the ramp. As indicated before, an additional major source of drag is the two counter-rotating vortex systems separating from the lower corner of the ramp, on its sides, grow in diameter, diffuse and eventually break as turning downstream. The flow field associated with the vortex system originating at the lower ramp corners could be seen in figures 20a and 20b. The flow structure obtained by both experiment and simulation is similar. A tight vortex separates from the lower ramp corner. It increases in diameter and is tilted in the downstream direction. This vortex system is unstable and has some common features with the vortex system over delta wings as well as the von-Karman vortex system behind circular cylinder. Its typical unsteadiness is at a Struhal number of order 0.2 based on the model width. Preliminary indications are that this vortex system is very receptive to excitation, as does the vortex system over a stalled delta wing (Ref. 1). However, this phenomenon was found to be sensitive to the symmetry of the vortex system. Therefore, comprehensive testing is necessary to first document and comprehend the structure of the baseline flow before attempting to control it. The findings will be presented in subsequent publications.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 13
Fig. 20a Computed Ramp vortex systems .
Fig. 20b Baseline Ramp right vortex system as seen by smoke flow visualization.
Conclusions, Recommendations and Future Research A combined experimental and numerical investigation was undertaken on a generic transport plane/helicopter fuselage in order to reduce the drag and alleviate unsteady loads resulting from the poor aerodynamics imposed by the presence of an aft loading ramp. The experiment included surface pressures, direct drag measurement and surface as well as smoke flow visualization. The numerical approach, applied at this stage only to the baseline flow was finite volume solutions of RANS equations, in steady and time-accurate modes. The baseline flow around the model was insensitive to the Reynolds number in the range it was tested. The flow separating from the aft body was characterized by two main sources of drag and unsteadiness. The first is a separation bubble residing at the lower ramp corner and the second is a pair of vortex systems developing and separating from the sides of the ramp. What seems to be a secondary bubble on the ramp causes increased suction and elevated drag as the model incidence is being reduced from positive to negative angles. As the incidence decreases the pair of vortex systems is penetrating deeper towards the centerline of the ramp. As expected, the ramp corner bubble was very receptive to periodic excitation introduced from the Piezo-fluidic actuators situated at the ramp lower corner. Total drag was reduced by 3-11%, depending on the model incidence. There are indications that the flow in the wake of the model is also significantly steadier when the bubble at the lower ramp corner is eliminated. The vortex system is tighter and steadier when the bubble is eliminated. Current tests include the effect of an additional actuator placed at the mid ramp location aimed at controlling the secondary ramp bubble. The effect of side-slip on the side vortices is being investigated prior to attempting to control it using the multiple actuators positioned at the sides of the ramp. Future tests should preferably include 3D PIV and a comprehensive comparison with numerical simulations, both steady and unsteady and eventually including the effects of the fluidic periodic excitation. To add validity to the current and future pressures and balance measured force data, a wind tunnel entry is planned to a large subsonic wind tunnel, eliminating tunnel interference and measuring all forces and moments directly. Closed-loop control would eventually be used for such complex application, especially when stabilizing the wake would remain the main issue. The next stages in the numerical analysis will be the use of a more appropriate turbulence model for separated flow analysis (v2f model) within the URANS analysis. This will be followed by simulations of flow control, using the above described approach and LES solutions.
AIAA 2004-2509
American Institute of Aeronautics and Astronautics 14
References
1. Margalit S, Greenblatt D, Seifert A, Wygnanski I., “Active flow control of a delta wing at high incidence using segmented piezoelectric actuators”. AIAA paper 2002-3270, June 2002.
2. Naim, A., Seifert, A. and Wygnanski, I., “Active Control of Cylinder Flow With and Without a Splitter Plate Using Piezoelectric Actuators”, AIAA Paper 2002-3070, June 02.
3. Arad, E., “Analysis of A Boundary Layer Separation over a Bump Using Large Eddy Simulation”, AIAA paper 2000-2558, 15th AIAA CFD conf., Anaheim, CA, 2001.
4. Seifert, A. and Pack, L.G., “Active Control of Separated Flow on a Wall-mounted “Hump” at High Reynolds Numbers”, AIAA J., V. 40, No. 7, July, 2002, pp. 1363-1372. (Part of AIAA paper 99-3430).
5. Arad E. and Martinelli L., “Large Eddy Simulation of Compressible Flow Using a Parallel, Multigrid Driven Algorithm”, AIAA-96-2065, 27th AIAA Fluid Dynamics Conf., New Orleans, LA, 1996.
![Page 1: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/1.jpg)
![Page 2: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/2.jpg)
![Page 3: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/3.jpg)
![Page 4: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/4.jpg)
![Page 5: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/5.jpg)
![Page 6: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/6.jpg)
![Page 7: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/7.jpg)
![Page 8: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/8.jpg)
![Page 9: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/9.jpg)
![Page 10: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/10.jpg)
![Page 11: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/11.jpg)
![Page 12: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/12.jpg)
![Page 13: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/13.jpg)
![Page 14: [American Institute of Aeronautics and Astronautics 2nd AIAA Flow Control Conference - Portland, Oregon ()] 2nd AIAA Flow Control Conference - Generic Transport Aft-Body Drag Reduction](https://reader042.documents.pub/reader042/viewer/2022020615/575095281a28abbf6bbf60e0/html5/page/14.jpg)
