Water Tunnel Flow Visualization: Insight into Complex Three-Dimensional Flowfields

Flow visualization experiments were performed on a series of generic fighter configurations at medium to high angles of attack in both a w ater tunnel and a w ind tunnel. The models were tested originally in a w ater tunnel to investigate methods for controlling the forebody and wing/LEX vortices for airframe control augmentation at high angles of attack. The same models were subsequently tested in a low-speed wind tunnel at the identical Reynolds number to compare the core locations and burst points of the LEX and wing vortices in the two test mediums. A u nique application of a smoke flow visualization method used primarily in low-speed wind tunnels for two-dimensional airfoil flow visualization was adapted to provide a direct means for comparison to the dye-flow results from the water tunnel. Direct comparisons of photographs from the two tests at various angles of attack on three different configurations showed nearly identical flow fields, confirming expectations that the test medium should have n o bearing on the flow field structure.

Voluminous research exists on the effect of combination of low Reynolds number flow over low aspect ratio flat plate wings for usage in micro-aerial vehicles and UAVs, as they also fly under similar conditions and combination of low aspect ratio and low Reynolds number. It has been well established by many researchers (Torres and Mueller in AIAA 42(5):865–873, 2004; Okamoto and Azuma in AIAA 49(6):1135–1150, 2011; Mizoguchi and Itoh in AIAA 51(7):7–52, 2013; Shields and Mohseni in AIAA 50(1):85–99, 2012; Mizoguchi and Kajikawa in Trans Jpn Soc Aero Sp Sci 59(2):56–63, 2016; Pelletier and Mueller in J Aircr 37(5):825–832, 2000; Liu and Hsiao in J Mech 28(01):77–89, 2012; Taira and Colonius in J Fluid Mech 623:187–207, 2009) that low aspect ratio wings in low Reynolds number flow gives a higher lift coefficient and stall angle is also increased which is quite different from the high aspect ratio wing. Also, it is a well-established fact in the existing literature that linear theories cannot be applied as in case of high aspect ratio wings because of non-linear increase in lift coefficient. The reason for such increase in lift and stall angle has been attributed to wingtip vortices that become prominent in low aspect ratio wings called the vortical lift which keeps on increasing up to a certain high angle of attack till it interacts with the separation bubble. In the past, very limited work has been done to study the vortices and their contribution in enhancement of aerodynamic characteristics. In the present work, flat plate model with corrugations at an aspect ratio (AR) of 1.0 are studied and compared with flat plate wing of AR 1.0; also flow visualization in water tunnel has been carried out at Reynolds number of 102 to ascertain the flow physics involved.

An experimental study of the breakdown of leading-edge vortices on diamond, cropped, delta, and double delta wings during dynamic pitching

A low-speed wind tunnel investigation was conducted to examine the vortex wake downstream of a three-percent scale model of the YF-17 lightweight fighter prototype at high angles of attack. The study was in support of NASA Ames Research Center's wind tunnel investigation of a full scale F/A-18 as part of NASA's High Alpha Technology Program. Smoke flow visualization was used to locate the downstream vortex wake. liotwire surveys were taken through the vortex at two stations: one directly aft of the model and the other at a station three model lengths downstream of the model. The effect of adding a fence to the leading edge extension (LEX) was studied. Power spectra from the hot-wire were recorded for the survey station directly aft of the model. Results show that peak turbulent fluctuation at this station occurred at 25 degrees angle of attack, lateral turbulent fluctuation greatly diminished at the far downstream station, and the addition of the LEX fence shifted energy content of turbulence toward higher frequencies.

Buffeting is an aeroelastic phenomenon that plagues high performance aircraft, especially those with twin vertical tails. Unsteady vortices emanate from wing/fuselage leading edge extensions when these aircraft maneuver at high angles of attack. These aircraft are designed such that the vortices shed while maneuvering at high angles of attack and improve the lift-to-drag ratio of the aircraft. With proper placement and sizing of the vertical tails, this improvement may be maintained without adverse effects to the tails. However, there are tail locations and angles of attack where these vortices burst and immerse the vertical tails in their wake inducing severe structural vibrations. The resulting buffet loads and severe vertical tail response became an airframe life and maintenance concern as life cycle costs increased. Several passive methods have been investigated to reduce the buffeting of these vertical tails with limited success. As demonstrated through analyses, wind-tunnel investigations, and full-scale ground tests, active control systems offer a promising solution to alleviate buffet induced strain and increase the fatigue life of vertical tails. A collaborative research project including the United States, Canada, and Australia is in place to demonstrate active buffet load alleviation systems on military aircraft. The present paper provides details on this collaborative project and other research efforts to reduce the buffeting response of vertical tails in fighter aircraft.

A comparative study of natural and ventilated supercavitation across two closed-wall water tunnel facilities

In support of the NASA High Alpha Technology Program, Navier-Stokes solutions have been obtained using the Chimera overset grid scheme for flow over the wing, fuselage, and wing leading-edge extension (LEX) of the F/A-18 High Alpha Research Vehicle (HARV) at high incidence. Solutions are also presented for flow over the fuselage forebody at high angles of attack. The solutions are for turbulent flows at high-Reynolds-number flight-test conditions, and are compared with available qualitative and quantitative experimental data. Comparisons of predicted surface flow patterns, off-surface flow visualization, and surface-pressure distributions are in good agreement with flight-test data. The ability of the numerical method to predict the bursting of the LEX vortex as it encounters the adverse pressure gradient field of the wing is demonstrated, and the capability of predicting high-angle-of-attack aerodynamics around realistic aircraft configurations is established.

Navier-Stokes solutions have been obtained using the Chimera overset grid scheme for flow over the wing, fuselage, and wing leading-edge extension (LEX) of the F-18 aircraft at high incidence. Solutions are also presented for flow over the fuselage forebody at high angles of attack. The solutions are for turbulent flows at high-Reynolds-number flight-test conditions, and are compared with available qualitative and quantitative experimental data. Comparisons of predicted surface flow patterns, off-surface flow visualizations, and surface- pressure distributions are in good agreement with flight-test data. The ability of the numerical method to predict the bursting of the LEX vortex as it encounters the adverse pressure gradient field of the wing is demonstrated. HIGH angle-of-attack technology program is currently underway Vvithin NASA. The objectives of the program include the development of flight-valida ted design methods that accurately predict the aerodynamics and flight dynamics of aircraft maneuvering in the high-angle-of-attack regime. Toward meeting these objectives, the program integrates ground-based experimental and computational investigations with flight-test investigations conducted on an F-18 at the NASA Ames-D den Flight Research Facility. The flight tests include surface 1 off-surface flow visualizations, as well as quantitative m irements of the flow surrounding the air- craft at large i ence. This database provides a unique op- portunity for C '•-).) code validation at actual high-alpha flight conditions. This paper presents results of Navier-Stokes com- putations of the flow about the wing, fuselage, and wing lead- ing-edge extension (LEX) of the F-18 at conditions matching those of the flight tests. Previous numerical predictions for the isolated F-18 fuselage forebody were reported in Refs. 1-3. Numerical prediction of the flow over aircraft flying at large angles of attack is a difficult aerodynamic problem. High- angle-of-attack flows contain large regions of three-dimen- sional separated flow, where the boundary layers leave the surface of the body along surfaces of separation, and roll up on the leeward side of the body to form strong, concentrated vortical flows. Separated flows historically have been treated by a wide variety of computational methods, ranging from simple potential flow methods to time-marching Navier-Stokes techniques. However, the close coupling that exists between the strength and location of the leeward vortical flow and the location of the viscous layer separation lines has precluded

Top inlet flow field and inlet performance data are presented which provide preliminary insight into the feasibility of upper-fuselage mounted inlet systems for transonic-supersonic fighter aircraft. Presented data span the Mach 0.2 to 2.0 envelope and enable evaluation of the influence of key aircraft configuration variables - inlet location, wing position, wing leading-edge extension (LEX) planform area, and variable incidence canards - on top inlet performance. The viability of this concept relative to more conventional inlet/airframe integrations is assessed via comparative evaluation of top and conventional inlet flow field parameters at transonic and supersonic speeds. It is shown that the action of the wing LEX vortex system produces a significant improvement in top inlet performance. Currently available transonic-supersonic data indicate that top inlet systems pose a viable configuration option for fighter aircraft requiring moderate angle of attack capability. However, recently acquired data indicate that increased angle of attack capability may be obtained by increasing wing leading-edge sweep angle.

Experimental study on pulsed suction under different momentum coefficients in variable compressor cascade with annular gaps

높은 받음각 상태에 있는 연장된 앞전을 갖는 편요된 삼각날개 주위의 와류 유동 특성에 대하여 날개 윗면에서의 압력 측정을 통하여 연구하였다. 본 연구에서는 와류 유동의 물리적 현상, 특히 높은 받음각 상태에서 받음각과 옆미끄럼각이 날개 형상 주위의 공력 특성에 미치는 영향에 중점을 두었다. 실험 데이터로부터 LEX 와류가 날개와류에 유익한 효과를 부여함을 명확하게 알 수 있었다. 이는, 작은 옆미끄럼각 상태에서는 비교적 높은 받음각에 이르기까지 와류의 붕괴 없이 날개와류를 안정화시킴을 알 수 있다. 특정 범위의 받음각과 옆미끄럼각 상태에서는 롤링 모멘트가 역전되는 흥미로운 유동 현상을 관찰할 수 있었다. An experimental study of the vortical flow characteristics around a yawed delta wing with the leading edge extension at high incidence angle is undertaken by upper surface pressure measurements. A special emphasis has been put on analyzing the basic physics of vortical flows, concerning the effects of incidence and sideslip angle on the aerodynamic characteristics of the wing, especially under high angle of attack. The experimental data has been dearly demonstrated the beneficial effect of the LEX vortex on the wing vortex. It leads to an essential stabilization of the wing vortex against its breakdown until at much higher incidence angle under small sideslip. An interesting flow feature is occurrence of the rolling moment reversal at a certain range of angle of attack and sideslip angle.

This paper embraces result of wind tunnel tests of a rocket plane designed to space tourism application. The rocket plane is designed in a tailless configuration with a leading edge extension (LEX) and side plates on the wing’s tip which work as all moving tail. Usually for such a concept of control, movable surfaces are position in a horizontal direction while in the considered concept there is no classical rudder, and the movable surfaces are mounted with a significant dihedral angle. The research was carried out in the subsonic closed circuit wind tunnel with an open test section. Moreover, MGAERO and PANUKL package were used for numerical computations. The first research question is to investigate how the configuration of the side plate affects the directional stability of the rocket plane. The second aim is to study the efficiency of the side plates deflected like an all moving tail for both low and high angles of attack. As a result of this investigation, the directional stability derivatives and the control derivatives were obtained. Finally, the experimental results were compared with numerical outcomes to establish does software with no full flow model can be applicable for design an aircraft with all moving tail in case of low of angles of attack.

Introduction T HE prediction of poststall maneuvers and departure characteristics of high-performance aircraft continues to be challenging,1;2 due, in particular, to the highly three-dimensional nature of the motions and aerodynamic responses. A case in point is the F/A-18 aircraft (Fig. 1), which has experienced departures from controlled  ight, typically involving large sideslip excursions with attendant high body-axes rate variations leading to high angles of attack ® (Ref. 4). The lateral–directional departure susceptibility of the F/A-18 provided part of the motivation within the NASA High-Alpha Technology Program for the improvement of high-alpha predictionsbased on ground testing. The gyrations and responsesexperiencedon departureare intrinsicallynonplanar, that is, involvingarbitraryrotationsnot alignedwith theprincipalaxesof the aircraft.3 When the  ight mechanical deŽ ciencies of an F/A-18 simulation were identiŽ ed, a Canadian programwas undertaken to enhance simulator Ž delity.6 The areas in which the airframe model could be improved included the poststall aerodynamicdatabase, the multivariabletable lookuparchitecture,and the aerodynamicmodel. The present paper focuses on the improvements to the database. The shortcomings of the poststall aerodynamic database were attributed largely to facility interference and scaling effects not accounted for and gaps in the parameter space covered. The original rotary database for F/A-18 was generated in the NASA Langley Research Center Spin Tunnel.11 The F/A-18 poststall maneuvering aerodynamics is dominated by forebody/leading edge extension (LEX) vortex interactions at moderately high angles of attack (30 55 deg). In the light of the sensitivity of the vortex interactions to sideslip angle (Ref. 12), it was deemed necessary to generate a more detailed rotary database to deŽ ne the associated relationshipsover a large sideslip range. When the advantages of the water-tunnel approachwere considered, it seemed appropriate to use this ground-testmethodology to generate supplementary data. The orbital platform rotary balance system (OPLEC) at the Institute for Aerospace Research (IAR) water tunnel, a diagnostic tool for studying support interference7;8 and unsteady wall interference, was subsequently used in highresolution tests as well as diagnostic experiments15i18 on the baseline F/A-18. A thorough understanding of scale effects on vortex

A wind tunnel investigation of the pressure field on the wing upper surface of a rigid 6% scale model of the F/A-18 was conducted in the Institute for Aerospace Research (IAR) 1.5m trisonic blowdown wind tunnel. The starboard wing and leading-edge extension were instrumented with 168 transducers for unsteady pressure measurements. The investigation was carried out for various Mach numbers and angles of attack. In this paper, detailed results at M = 0.6 and a = 30' with a clean wing configuration are presented. The emphasis of this study is on the structure of the f lowfield on the wing upper surface at high angles of attack. Time-averaged and rms pressure distributions were determined and the normal force on the wing was computed for different angles of attack. Spectral analyses of the pressures at different locations on the wing were carried out. Broadband spacetime correlation of the pressure transducer signals was performed and the convection pattern of the pressure field was compared to the direction of the skin friction lines obtained from flow visualization studies using oil dots deposited on the wing surface.

Due to flow asymmetries created in the tip region, aircraft forebodies can experience large net side forces at high angles of attack. It has been shown that these forces can be controlled by the use of suction in the tip region. For this purpose, a control system that manipulated the flow near the tip was tested on an aircraft forebody model in a wind tunnel. Utilizing a conventional PD algorithm, the system was able to control the forebody satisfactorily when tuned for a particular Reynolds number; but the performance degraded as the Reynolds number was changed. Fuzzy logic control (FLC) was then applied and shown to be more robust than PD. However, the FLC also showed some degradation for flows at different Reynolds numbers. To address this problem, an adaptive fuzzy controller was used. This paper discusses fuzzy model reference learning control (FMRLC) as applied to the aircraft model. It is shown that FMRLC is able to quickly and automatically generate a rule base to control the model. It is then shown that he FMRLC is able to modify it’s rule base and adapt in real time when the Reynolds number of the system changes.