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Experimental Investigation of Actuators for Flow Control in Inlet Ducts
Abstract
Attractive to aircraft designers are compact inlets, which implement curved flow paths to the compressor face. These curved flow paths could be employed for multiple reasons. One of which is to connect the air intake to the engine embedded in the aircraft body. A compromise must be made between the compactness of the inlet and its aerodynamic performance. The aerodynamic purpose of inlets is to decelerate the oncoming flow before reaching the engine while minimizing total pressure loss, unsteadiness and distortion. Low length-to-diameter ratio inlets have a high degree of curvature, which inevitably causes flow separation and secondary flows. Currently, the length of the propulsion system is constraining the overall size of Unmanned Air Vehicles (UAVs), thus, smaller more efficient aircrafts could be realized if the propulsion system could be shortened. Therefore, active flow control is studied in a compact (L/D=1.5) inlet to improve performance metrics. Actuation from a spanwise varying coanda type ejector actuator and a hybrid coanda type ejector / vortex generator jet actuator is investigated. Special attention will be given to the pressure recovery at the AIP along with unsteady pressure signatures along the inlet surface and at the AIP.
... Owens et al. (2008) installed a similar single row of vortex generators into an s-duct operating in a free stream at M = 0.85, using a model that more closely resembles the blended wing body flush-mounted s-duct. Boundary layer ingestion and the control of the resulting flow field using similar passive flow control devices was studied by Anabtawi et al. (1999) and Tournier and Paduano (2005). The induced drag associated with vanes was decreased through the use of sub-boundary layer devices (Anderson et al. 2002). ...
... A variety of locations and configurations of different flow injection were tested with resulting insight into the most effective and most efficient implementation for these particular jets. Vaccaro et al. (2010) showed that the use of Coanda surfaces along with various spanwise continuous jet configurations was effective at controlling the pressure recovery and reducing the intensity of the structures produced by the separation found in a rectangular offset duct. The parasitic use of bleed air was demonstrated to be avoidable for active flow control in ducts at low speeds by Amitay et al. (2008) who successfully used synthetic jets, which require zero mass flux, to control flow separation in an offset duct. ...
The formation and interaction of streamwise vorticity generated by surface-mounted passive (micro-vanes) and active (synthetic jets) flow control elements is investigated experimentally in a small-scale wind tunnel at high subsonic speeds (M < 0.6). Streamwise vortices are generated within the boundary layer of a converging-diverging wall insert that is designed to provide an adverse pressure gradient similar to the pressure gradient within a typical offset diffuser. Hybrid actuation is demonstrated by combining the micro-vanes with synthetic jets in tandem or by integrating the jets into vanes. Tandem actuation is accomplished using synthetic jet actuators that are yawed to match the vanes' yaw angle and skewed to generate streamwise vortices of matching sense. Both approaches lead to augmentation of the primary vanes' vortices and enhance their pairing. These approaches provide "fail-safe" control devices having robust nominal control effectiveness by the vane elements coupled with enhancement by the jets.
... They suppressed flow separation successfully, but also lead to an increased swirl in the flow at the AIP. As stated by Vaccaro et al. (2010), a 2D control jet actuator in an s-duct (Ma = 0.43) can affect centreline reattachment, but cannot influence the formation of secondary flows, and does not reduce pressure losses at Fig. 11. Normalised total pressure recovery contours for convoluted duct without (left) and with (right) vortex generator array. ...
This paper is a review of significant studies in the complex flow physics in diffusive, s-shaped ducts, focusing on flow control methods employed to counteract the onset of separation, swirl formation, and non-uniformity of pressure at the duct exit plane. Passive, active, and hybrid flow control, along with optimisation techniques used to control the dominant flow features are discussed. According to the literature, tapered fin vortex generators and submerged vortex generators improve pressure loss and distortion by double digit percentages, and three-dimensional synthetic jets and pulsed micro-jets show greatest promise amongst active flow control devices. Plasma flow control methods have only sparsely been used in s-ducts with one study performing experiments with alternating-current dielectric-barrier-discharge plasma actuators. The importance of flow unsteadiness has been identified in the literature, with peak values as high as one order of magnitude different from the time-averaged properties. Despite this, very few flow control studies have used time-dependent solution methods to quantify the effect of flow control methods on the unsteadiness of the flow.
In the current work, computational study has been carrying out on the both baseline serpentine duct to know the flow behaviour and baseline with vortex generator to know the effect of vortex generators on the flow properties of the duct. A set of vortex generators are chosen for the study, by keeping all other parameters constant. Here in current work, for modelling the duct and vortex generator CATIA V5 is used; to discretizing the geometry ICEM CFD is used. For the analysis, the computational fluid dynamics tool FLUENT is used, SST k-ù model available in the FLUENT is used for the computations, and the results of current work will be validated with experimental results and shows that they are converged. Use of vortex generator in the current work proved extremely effective in enforcing the active flow control. Results obtained in this work are to study the pressure recovery at the engine face. The results ensure that the current work on the baseline model of serpentine inlet duct and model with vortex generator comparison is well established in case of pressure recovery and decrease in engine face distortion. After comparing the result of both baseline duct and duct with vortex generator we are expecting at least 25-35% of pressure recovery at the engine face. In the present study significant improvements characterized by an improved pressure recovery and decrease in distortion are observed over the baseline model of serpentine duct for vortex generator model. This reduction in pressure loss and distortion enhances the engine performance of thrust and also improves the overall performance of the engine.
A combined experimental and numerical investigation of flow control actuation via a steady two-dimensional tangential control jet in a short inlet duct was conducted. Experiments were run on an inlet with a length to diameter ratio of 1.5 at a Mach number of 0.43 at both baseline and forced conditions. The setup in the numerical simulations matched the experimental tests in terms of both the physical parameters (e.g., Mach number, pressure level, etc.) and dimensions (e.g., contraction, inlet duct, control jet, etc.). Both experimental and numerical results for the base flow showed a quasi two-dimensional flow separation immediately downstream of the first (or upstream) turn on the lower surface. For the forced case it was found that actuation via a steady two-dimensional jet was able to control flow separation in the mid-span portion of the duct but was not effective towards the outer span and the side walls. The control jet increased the three-dimensionality of the flow field inside the inlet. In both cases, simulations were found to over predict separations found within the duct. This over prediction led to quantitative disagreement, nevertheless prominent flow structures were found to be qualitatively similar, and therefore, details of the three-dimensional flow field were explored based on numerical data. These structures were measured experimentally by means of pressure recovery, surface pressure distribution on the duct's lower wall and by the volumetric flow field, which was acquired using stereoscopic PIV (SPIV). Both experimental and numerical results infer that a two-dimensional control jet induces three-dimensional structures that are considerably larger than the ones found in the baseline flow field. © 2010 by the American Institute of Aeronautics and Astronautics, Inc.
It was recently demonstrated that oscillatory blowing can delay separation from a symmetrical airfoil much more effectively than the steady blowing used traditionally for this purpose, Experiments carried out on different airfoils revealed that this flow depends on many parameters such as, the location of the blowing slot, the steady and oscillatory momentum coefficients of the jet, the frequency of imposed oscillations, and the shape and incidence of the particular airfoil, In airfoils equipped with slotted flaps, the flow is also dependent on the geometry of the slot and on the Reynolds number In addition to the flap deflection that is considered as a part of the airfoil shape, The incremental improvements in single element airfoil characteristics are generally insensitive to a change in Reynolds number, provided the latter is sufficiently large, The imposed oscillations do not generate large oscillatory lift nor do they cause a periodic meander of the c.p.
The effects of oscillatory blowing as a means of delaying separation are discussed. Experiments were carried out on a hollow, napped NACA 0015 airfoil equipped with a two-dimensional slot over the hinge of the flap. The flap extended over 25% of the chord and was deflected at angles as high as 40 deg. The steady blowing momentum coefficients could be varied independently of the amplitudes and frequencies of the superimposed oscillations. The modulated blowing was a major factor in improving the performance of the airfoil at much lower energy inputs than was hitherto known. Optimum benefits in performance were obtained at reduced frequencies, based on the flap chord, of an order of unity. Significant increase in lift as well as cancellation of form drag were observed. The increase in Reynolds number did not have an adverse effect on the data.
Vortex generating jets (VGJs) are jets that pass through a wall and into a crossflow to create a dominant streamwise vortex that remains embedded in the boundary layer over the wall. The VGJ is characterized by its pitch and skew angles (Φ and Θ) and the velocity ratio (VR) between the jet and the crossflow. For VR=1.0, the VGJ configuration of Φ=30°,Θ=60° has been identified as that which produces the vortex with the highest peak mean vorticity. Three-component laser Doppler velocimetry (LDV) data for this particular configuration demonstrate many interesting features of the flow. Mean velocity data show a deficit of streamwise momentum in the core of the vortex, thinning of the boundary layer on the downwash side of the vortex, and thickening of the boundary layer on the upwash side. Plots of the turbulent kinetic energy and the turbulent shear stress 〈uv〉 show that the turbulent structure of the boundary layer is grossly disturbed by the presence of the vortex. The turbulent transport of the turbulent kinetic energy shows the possibility for a gradient diffusion model in most regions, but not the vortex core.
This paper presents results from a joint Lockheed Martin/NASA Glenn effort to design and verify an ultra-compact, highly-survivable engine inlet subsonic duct based on the emerging technology of active inlet flow control (AIFC). In the AIFC concept, micro-scale actuation (approx. mm in size) is used in an approach denoted 'secondary flow control' to intelligently alter a serpentine duct's inherent secondary flow characteristics with the goal of simultaneously improving the critical system-level performance metrics of total pressure recovery, spatial distortion, and RMS turbulence. In this approach, separation control is a secondary benefit, not a design requirement. The baseline concept for this study was a 4:1 aspect ratio ultra-compact (L/D = 2.5) serpentine duct that fully obscured line-of-sight view of the engine face. At relevant flow conditions, this type of duct exhibits excessive pressure loss and distortion because of extreme wall curvature. Two sets of flow control effectors were designed with the intent of establishing high performance levels to the baseline duct. The first set used two arrays of 36 co-rotating microvane vortex generators (VGs); the second set used two arrays of 36 micro air-jet (microjet) VGs, which were designed to produce the same 'vorticity signature' as the microvanes. Optimisation of the microvance array was accomplished using a design of experiments (DOE) methodology to guide selection of parameters used in multiple Computational Fluid Dynamics (CFD) flow solutions. A verification test conducted in the NASA Glenn W1B test facility indicated low pressure recovery and high distortion for the baseline duct without flow control. With microvane flow control, at a throat Mach number of 0.60, pressure recovery was increased 5%, and both spatial distortion and turbulence were decreased approximately 50%. Microjet effectors also provided significantly improved performance over the baseline configuration.
The present paper discusses the effect of active flow control, via steady and unsteady control jets, on the performance of a very aggressive (length to exit diameter ratio, L/D, of 1.5) inlet duct. The experiments were performed at an inlet Mach number of 0.45 in a specially designed facility, which enables testing of various actuation methodologies. Two major parameters were compared: the effect of blowing ratio (at a fixed momentum coefficient), and the actuation frequency. Although the facility is equipped with a pair of control jets (to mitigate separation on the two duct bends) only the upstream control jet was used to simplify the experiments. Steady and unsteady pressure measurements (along the lower wall of the duct, at the AIP, and in the actuator's cavity) were performed for various momentum coefficients, blowing ratios, and actuation frequencies. Flow control was shown to have a substantial effect, mainly along the centerline. The data suggest that using 2-D flow control to affect a flow field that is highly three-dimensional is not optimal, and as such a spanwise varying actuation should be implemented. In addition to the experimental effort, initial results on the optimal sensor placement and estimation of pressure recovery over a portion of the AIP are presented. Data from a two-dimensional CFD simulation is used to generate flow field data for a throat Mach number of 0.45 and several blowing input profiles on the lower surface. Results are presented for AIP estimation using both static and dynamic models, and it is shown that acceptable performance can be obtained for all values of blowing input considered here. Static and dynamic models can predict average lower AIP pressure recovery with RMS error of .05% and worst-case error of about 0.6%. Sensor placement analysis is performed by iteratively testing combinations of candidate sensor locations and tabulating them that most often are part of the combinations with least error. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.
S-duct Inlets are commonly used on subsonic cruise missiles, as they offer a good compromise between compactness, low observability and aerodynamic performance. Though currently used S-ducts exhibit good performance in terms of distortion and pressure recovery at the AIP, the situation can degrade drastically when the inlet is put in off-design conditions, with the risk of compressor instabilities. Flow control is considered as a promising way to maintain inlet efficiency in off-design flight conditions. Industrial interest for flow control techniques is therefore rising, and a need for their comparative evaluation has been expressed. In response to this need, an experimental setup has been designed and fabricated, and flow control experiments have been carried out at MIT, on the selected off-design case of forebody boundary layer ingestion. The first set of experiments focused on the characterization of the inlet in a clean configuration. Then, a distortion device was added in order to simulate thick forebody boundary layer. The introduced disturbance proved to have a strong detrimental effect on the inlet performance, as the separation bubble grew in size, the pressure recovery dropped down and the distortion level Increased drastically. The selected flow control techniques were then implemented. Initial results were discussed in a previous paper 14. Vortex Generators in different geometrical configurations improved the pressure recovery and significantly decreased the distortion level. Their use had a strong impact on the flow structure, delaying or even suppressing separation. Steady injection was implemented with Coanda-type injectors, upstream of the separation line. It led to significant Improvement of the pressure recovery, which increased with increasing injection mass flow. Separation was eliminated for the highest injection mass flow. The distortion level decreased with increasing injection mass flow. Overall, the results also highlighted the importance of the secondary flows as a source of distortion and pressure recovery loss.
A series of flow control experiments were carried out in order to reconcile some of the observations made about enhancement of lift by periodic excitation. The relative merits of flow control were assessed in comparison with conventional design techniques to establish the best strategy to utilize flow control. The frequency of choice, its effect on the pressure distribution, and the optimum location of the actuation and its amplitudes are specifically discussed. The results of periodic excitation were also compared to results generated using steady blowing or suction
Active flow control, via steady control jets, was implemented to improve the performance of a very aggressive (length to exit diameter ratio, L/D, of 1.5) inlet duct. The experiments were performed for a range of inlet Mach number from 0.2 to 0.45. A brand new facility was designed and built to enable various actuation methodologies as well as multiple measurement techniques. In the present work, a pair of steady control jets was placed in streamwise locations where flow was expected to separate. Static pressure measurements, along the upper and lower walls of the duct, were performed for various combinations of actuation. The forcing level of the control jets as well as combinations of jets, were tested. In addition, total pressure measurements were conducted at the Aerodynamic Interface Plane (AIP) to obtain the distribution of the pressure recovery. Flow control was shown to have a substantial effect, mainly on the lower wall. It was found to be more effective at the lower Mach number where the blowing ratio was higher (for the same mass flux ratio). The data suggest that using 2-D flow control to affect a flow field that is highly three-dimensional is not optimal, and as such a spanwise varying actuation should be implemented. Copyright © 2008 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
The effect of active flow control, via arrays of synthetic jet actuators, on the flow field around a finite wing was investigated experimentally and numerically. To fully and properly implement flow control, a fundamental understanding of the interaction of the synthetic jets with the three-dimensional cross flow must first be possessed. The experiments were conducted in a wind tunnel on a finite wing having a cross-sectional profile of NACA 4421 at a wide range of angles of attack, Reynolds numbers, and several arrangements of synthetic jets. The majority of the current work has been focused on the interaction of the jets with the cross-flow at low angles of attack (where separation is not present). Two dimensional and stereoscopic PIV data were collected in conjunction with dynamic surface pressure. For the numerical investigation, a finite-span synthetic jet surface mounted on a NACA 4421 airfoil was analyzed using an in house flow solver (PHASTA). Using an adaptive meshing technique on an unstructured grid and Streamline Upwind Petrov-Galerkin stabilized finite element formulation, the behavior of the synthetic jets in a cross flow was simulated in three dimensions. Using these data, the complex 3-D interactions were analyzed to form a cohesive understanding of the parameters that may impact the effectiveness of flow control on 3-D configurations. Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc.
The effect of the actuation frequency on the manipulation of the global aerodynamic forces on lifting surfaces using surface-mounted fluidic actuators based on synthetic (zero mass flux) jet technology is demonstrated in wind-tunnel experiments. The effect of the actuation is investigated at two ranges of (dimensionless) jet formation frequencies of the order of, or well above, the natural shedding frequency. The vortical structures within the separated flow region vary substantially when the dimensionless actuation frequency F+ is varied between O(1) and O(10). When F+ is O(1), the reattachment is characterized by the formation of large vortical structures at the driving frequency that persist well beyond the trailing edge of the airfoil. The formation and shedding of these vortices leads to unsteady attachment and, consequently, to a time-periodic variation in vorticity flux and in circulation. Actuation at F+ of O(10) leads to a complete flow reattachment that is marked by the absence of organized vortical structures along the flow surface. This suggests that when the actuation frequency is high enough, the Coanda-like attachment of the separated shear layer to the top (suction) surface of the airfoil can be replaced by completely attached flow for which separation may be bypassed altogether.
The objectives of the study described here are twofold: (1) to
demonstrate experimentally if acoustic excitation can control flow
separation over airfoils of different sizes, and establish scaling
between the acoustic parameters, airfoil size, and flow velocity; (2) to
evaluate experimentally if sound emitted within the airfoil body is as
effective as sound irradiated from an external source in controlling
flow separation. These objectives have been accomplished by conducting
model-scale experiments in high-quality wind tunnels using existing
instrumented airfoils. Both quantitative data and flow visualization
results are presented.
The development of free shear layers formed by the mixing of initially separated free streams is examined in a review of recent work. The mixing layer is viewed as a prototype for a class of inviscidly unstable free shear flows including jets and wakes, and the focus is on 2D homogeneous incompressible mixing layers. Major areas covered include dynamical processes in free shear layers, the influence of operational parameters, sensitivity to artificial excitations, and global feedback effects. Graphs, drawings, and photographs of characteristic phenomena are provided.
The effect of periodic two-dimensional excitation on the development of a turbulent mixing region was studied experimentally. Controlled oscillations of variable ampli- tude and frequency were applied at the initiation of mixing between two parallel air streams. The frequency of forcing was at least an order of magnitude lower than the initial instability frequency of the flow in order to test its effect far downstream. The effect of the velocity difference between the streams was also investigated in this experiment. A typical Reynolds number based on the velocity difference and the momentum thickness of the shear layer was l0 ⁴ .
It was determined that the spreading rate of the mixing layer is sensitive to periodic surging even if the latter is so small that it does not contribute to the initial energy of the fluctuations. Oscillations at very small amplitudes tend to increase the spreading rate of the flow by enhancing the amalgamation of neighbouring eddies, but at higher amplitudes the flow resonates with the imposed oscillation. The resonance region can extend over a significant fraction of the test section depending on the Strouhal number and a dimensionless velocity-difference parameter. The flow in the resonance region consists of a single array of large, quasi-two-dimensional vortex lumps, which do not interact with one another. The exponential shape of the mean-velocity distribution is not affected in this region, but the spreading rate of the flow with increasing distance downstream is inhibited. The Reynolds stress in this region changes sign, indicating that energy is extracted from the turbulence to the mean motion; the intensity of the spanwise fluctuations is also reduced, suggesting that the flow tends to become more two-dimensional.
Amalgamation of large coherent eddies is resumed beyond the resonance region, but the flow is not universally similar. There are many indications suggesting that the large eddies in the turbulent mixing layer at fairly large Re are governed by an inviscid instability.
The phenomenon of flow attachment to solid surfaces, occurring with both liquids and gases, is the long-known though inadequately understood Coanda effect.
A flow visualization study was made using a birefringent milling yellow dye solutions flowing over a deflection surface consisting of flat plates. A two-dimensional flow channel with transparent side walls was used. Photographic observations of the development of the Coanda effect reveal the method of flow attachment and confirm a number of literature predictions. One of the most interesting of these phenomena is the existence of a well defined mixing region along the deflection surface.
A simplified model of the flow field has been proposed in order to described the mechanism governing the Coanda effect. The model is supported by experimental data consisting of pressure profiles obtained along the deflection surface and secondary flow entrainment measurements.
The suppression of post-stall separation over an unconventional 2-D airfoil at moderate Reynolds numbers (up to 106) using synthetic (zero net mass flux) jet actuators is discussed. As shown by the authors in earlier investigations, the
apparent modification of the surface shape by the interaction domain between the actuator jets and the cross flow results
in a local displacement of the cross flow streamlines. The concomitant modification of the streamwise pressure gradient upstream
of where the flow nominally separates in the baseline configuration can lead to complete suppression of separation over a
significant range of angles of attack in the post-stall domain. While in the absence of flow control the airfoil is stalled
at angles of attack exceeding 5°, actuation leads to either completely or partially attached flow within the entire range
of angles tested (up to 25°) that is accompanied by a dramatic increase in lift and a corresponding decrease in pressure drag.
Actuation is typically effected at frequencies that are an order of magnitude higher than the characteristic (shedding) frequency
of the airfoil [i.e., F+ ∼ O(10) rather than F+ ∼ O(1)]. When the actuation frequency F+ is O(1), the reattachment is characterized by a Coanda-like tilting of the separated shear layer and the formation of large
vortical structures at the driving frequency that persist beyond the trailing edge of the airfoil and lead to unsteady attachment
and consequently to a time-periodic variation in vorticity flux and in circulation. In contrast, the suppression of separation
at high actuation frequencies [i.e., F+=O(10)] is marked by the absence of organized vortical structures along the flow surface. The dynamics of the transient lift
in controlled reattachment and separation are investigated using pulsed amplitude modulation of the actuation input and is
exploited to improve the efficacy of the jet actuators by using pulse modulation of the excitation input.
S-duct inlets are commonly used on subsonic cruise missiles, as they offer a good compromise between compactness, low observability and aerodynamic performance. Though currently used S-ducts exhibit good performance in terms of distortion and pressure recovery at the AIP, the situation can degrade drastically when the inlet is put in off-design conditions, with the risk of compressor instabilities. Flow control is considered as a promising way to maintain inlet efficiency in off-design flight conditions. Industrial interest for flow control techniques is therefore rising, and a need for their comparative evaluation has been expressed. In response to this need, an experimental setup has been designed and fabricated, and flow control experiments have been carried out at MIT, on the selected off-design case of forebody boundary layer ingestion. The first set of experiments focused on the characterization of the inlet in a clean configuration. Then, a distortion device was added in order to simulate thick forebody boundary layer. This proved to have a strong detrimental effect on the inlet performance, as the separation bubble grew in size, the pressure recovery dropped down and the distortion level increased drastically. The selected flow control techniques were then implemented. .
Numerical simulations of a single low-profile vortex generator vane, which is only a small fraction of the boundary-layer thickness, and a vortex generating jet have been performed for flows over a flat plate. The numerical simulations were computed by solving the steady-state solution to the Reynolds-averaged Navier-Stokes equations. The vortex generating vane results were evaluated by comparing the strength and trajectory of the streamwise vortex to experimental particle image velocimetry measurements. From the numerical simulations of the vane case, it was observed that the Shear-Stress Transport (SST) turbulence model resulted in a better prediction of the streamwise peak vorticity and trajectory when compared to the Spalart-Allmaras (SA) turbulence model. It is shown in this investigation that the estimation of the turbulent eddy viscosity near the vortex core, for both the vane and jet simulations, was higher for the SA model when compared to the SST model. Even though the numerical simulations of the vortex generating vane were able to predict the trajectory of the stream-wise vortex, the initial magnitude and decay of the peak streamwise vorticity were significantly under predicted. A comparison of the positive circulation associated with the streamwise vortex showed that while the numerical simulations produced a more diffused vortex, the vortex strength compared very well to the experimental observations. A grid resolution study for the vortex generating vane was also performed showing that the diffusion of the vortex was not a result of insufficient grid resolution. Comparisons were also made between a fully modeled trapezoidal vane with finite thickness to a simply modeled rectangular thin vane. The comparisons showed that the simply modeled rectangular vane produced a streamwise vortex which had a strength and trajectory very similar to the fully modeled trapezoidal vane.
Synthetic Jets Active Inlet Flow Control Technology Demonstration Perturbed Free Shear Layers
- A Glezer
- M Amitay
- J W Hamstra
- D N Miller
- P P Truax
- B A Anderson
- B Wendt
Glezer, A. and Amitay, M., " Synthetic Jets ", Annual Review of Fluid Mechanics, 2002, 34. Hamstra, J. W., Miller, D. N., Truax, P. P., Anderson, B. A. and Wendt, B., J., " Active Inlet Flow Control Technology Demonstration ", The Aeronautical Journal, Oct. 2000, pp. 473. Ho., C.-M. and Huerre, P., 1984, " Perturbed Free Shear Layers ", Annual Review of Fluid Mechanics, Vol. 16, pp. 365-424.
The use of a vibrating ribbon to delay separation on two dimensional airfoils Proceedings of Air Force Academy Workshop in Unsteady Separated Flow 1987 The Forced Mixing Layer Between Parallel Streams
- D Neuberger
- I Wygnanski
Neuberger, D. and Wygnanski, I., " The use of a vibrating ribbon to delay separation on two dimensional airfoils ", Proceedings of Air Force Academy Workshop in Unsteady Separated Flow 1987. Colorado Springs, CO. Oster, D. and Wygnanski, I.J., 1982, " The Forced Mixing Layer Between Parallel Streams ", J. Fluid Mech., 123, pp. 91-130.
FlowAnalysisand Control ina Subsonic Inlet
- S Tournier