Chord Airfoil Research Articles

Aerodynamic Characteristics of Morphing Supercritical Airfoils for Aircraft with All-Stage High Performance

Morphing airfoil is a promising technology for future aircraft to realize all-stage high performance. In the present paper, a conceptual aircraft with morphing airfoil is proposed and the aerodynamic characteristics of three types of morphing airfoils (variable-camber airfoil, variable-chord airfoil, and the combination of both morphing styles) are numerically investigated. The baseline airfoil is RAE 2822 supercritical airfoil; the Reynolds-averaged Navier–Stokes method is adopted for numerical simulation of flow around airfoils, and the accuracy of the numerical simulation method is validated by comparing with experimental data. It is found that the variable-camber and -chord airfoil can not only improve the high lift characteristics at take-off stage, but also increase the lift-to-drag ratio at transonic cruise and low-speed task stages during which the required lift is continuously decreasing due to the consumption of fuel. These findings imply that aircraft with proper morphing airfoil can achieve all-stage high aerodynamic performance.

Numerical Investigation on Aerodynamic Characteristics of Damaged Infinite Wings With Variation in Penetration Angle

Aerodynamic performance of a full span NACA 641-412 airfoil with a circular-shaped damage at various attack directions has been numerically investigated in this study. To assess the aerodynamic effects of different penetration angles in which threats such as projectiles can pass through the wings, attack directions of 30°, 60°, -30° and -60° relative to the normal axis of the chord line has been studied and compared with attack direction of 0°. To validate with published studies about damaged wing, the 200 mm chord airfoil was simulated with the damage hole diameter of 20% chord at the midspan and midchord location in Reynolds number of 500,000. Quantitative and qualitative results of this numerical study had a good agreement with published experimental data due to appropriate structured mesh and turbulence modelling. In addition to lift, drag and pitching moment coefficient, surface pressure distribution around the damage hole has been studied. Results show that, if the penetration angle becomes more negative, aerodynamics performance of the wing will be further decreased; therefore, attack directions of threat mechanisms such as “ahead and above” or “below from the rear” have severe negative impact than other directions on aerodynamic performance of the damaged infinite wing.

Droplet Deformation and Trajectory Without Interferences in the Incoming Airfoil Shoulder Region

A new experimental campaign has been conducted in the shoulder region of an incoming airfoil of chord 1050 mm, where droplets are separated enough to neglect interferences between them. Droplets of three sizes (500, 950, and of radius) were allowed to fall in the path of an incoming airfoil while shadowgraph images were recorded by a high-speed video camera at 40,000 fps. The airfoil model was placed at the end of a rotating arm and moved at four velocities (30, 40, 50, and ). Three different regions of the shoulder were tested. Droplet deformation and trajectories are presented. Droplets evolve as a conjunction of two half-oblate spheroids that tilt as the model approaches. The tilting is larger in the higher regions of the shoulder. The trajectory model derived for droplet in the stagnation line of a moving airfoil has been formally derived for the shoulder region and applied to the experimental data, showing very good agreement being the mean discrepancy less than 4% for the trajectory and 10% for the deformation.

Investigation of full and partial ground effects on a flapping foil hovering above a finite-sized platform

The full and partial ground effects on the lift generation of a flapping airfoil in normal hovering mode are investigated numerically using the discrete vortex method in two dimensions. To achieve full ground effect, the airfoil of chord c is made to hover above the center of a finite-sized platform of length 10c. We have observed the force-enhancement, force-reduction, and force-recovery regimes at low, medium, and high ground clearances in line with the existing literature. This paper puts special focus on partial ground effect when the airfoil is hovering near the edge of the platform. Lift-modifying mechanisms not previously observed under full ground effect have been discovered. When stroke reversal occurs near the edge of the platform, a relatively stationary strong vortex may form above the platform edge. This strong vortex can either increase or decrease the instantaneous lift force on the airfoil depending on the position of the airfoil relative to the platform edge. Also, the platform edge may lead to the formation of an additional vortex pair which increases the instantaneous lift force as the airfoil sweeps past the edge under suitable conditions. Lastly, the platform edge can lead to the formation of a reverse von Kármán vortex street that extends well below the stroke plane under suitable geometric arrangements.

Modeling the increase in aerodynamic efficiency for a thick (37.5% chord) airfoil with slot suction in vortex cells with allowance for the compressibility effect

The Reynolds equations closed using the Menter shear-stress-transfer model modified with allowance for the curvature of flow lines have been numerically solved using multiblock computational technologies. The obtained solution has been used to analyze subsonic flow past a thick (37.5% chord) airfoil with slot suction in circular vortex cells intended for the Ecology and Progress (Ekologiya i Progress, EKIP) aircraft project in comparison to a distributed (from the central body surface) suction at fixed values of the total volume flow rate (0.02121) and Reynolds number (105) in a range of Mach numbers from 0 to 0.4. This analysis revealed a significant (up to tenfold) decrease in the bow drag (determined with allowance for the energy losses) and a large (by an order of magnitude) increase in the aerodynamic efficiency of the thick airfoil containing vortex cells with slot suction, which reached up to 160.

Computational Analysis of Propulsion Performance of Modified Pitching Motion Airfoils in Laminar Flow

The thrust generation performance of airfoils with modified pitching motion was investigated by computational fluid dynamics (CFD) modeling two-dimensional laminar flow at Reynolds number of 104. The effect of shift distance of the pitch axis outside the chord line(R), reduced frequency(k), pitching amplitude(θ), pitching profile, and airfoil shape (airfoil thickness and camber) on the thrust generated and efficiency were studied. The results reveal that the increase inRandkleads to an enhancement in thrust generation and a decrease in propulsive efficiency. Besides, there exists an optimal range ofθfor the maximum thrust and the increasingθinduces a rapid decrease in propulsive efficiency. Six adjustable parameters(K)were employed to realize various nonsinusoidal pitching profiles. An increase inKresults in more thrust generated at the cost of decreased propulsive efficiency. The investigation of the airfoil shape effect reveals that there exists an optimal range of airfoil thickness for the best propulsion performance and that the vortex structure is strongly influenced by the airfoil thickness, while varying the camber or camber location of airfoil sections offers no benefit in thrust generation over symmetric airfoil sections.

Propagating-arc magnetohydrodynamic plasma actuator for directional high-authority flow control in atmospheric air

A propagating-arc magnetohydrodynamic plasma actuator for aerodynamic flow control is reported. The actuator comprises two rail electrodes flush mounted on an aerodynamic surface. A pulsed arc is propelled down the length of the rails by Lorentz forces supported by a self-induced magnetic field. The arc induces a high velocity pulsed air wall jet due to the pushing and entrainment actions. Experiments in quiescent air demonstrate that the plasma arc achieves a peak velocity of around 100 m s−1 and requires a discharge energy on the order of 300 J per pulse. Wind tunnel tests on a 14.5 inch chord airfoil section, at a Reynolds number of 0.45 million show induced flow velocities on the order of 10's m s−1 with significant penetration of the flow actuation effect perpendicular to the wall surface.

Novel, Bidirectional, Variable-Camber Airfoil via Macro-Fiber Composite Actuators

This study aims to enable solid-state aerodynamic force generation in high-dynamic-pressure airflow. A novel, high-load-output, bidirectional variable-camber airfoil employing a type of piezoceramic composite actuator known as a Macro-Fiber Composite is presented. The novel airfoil employs two active surfaces and a single four-bar (box) mechanism as the internal structure. The unique choice of boundary conditions allows variable and smooth deformation in both directions from a flat camber line. The paper focuses on actuation modeling and response characterization under aerodynamic loads. A parametric study of aerodynamic response is employed to optimize the kinematic parameters of the airfoil. The concept is fabricated by implementing eight Macro-Fiber Composite 8557-P1-type actuators in a bimorph configuration to construct the active surfaces. The box mechanism generates deflection and camber change as predicted. Wind-tunnel experiments are conducted on a 12.6% maximum thickness, 127 mm chord airfoil. Aerodynamic and structural performance results are presented for a flow rate of 15 m /s and a Reynolds number of 127,000. Nonlinear effects due to aerodynamic and piezoceramic hysteresis are identified and discussed. A lift coefficient change of 1.54 is observed purely due to voltage actuation. Results are compared with conventional, zero-camber NACA and other airfoils. A 72% increase in the lift-curve slope is achieved when compared with a NACA 0009 airfoil.

Edge-source acoustic green's function for an airfoil of arbitrary chord, with application to trailing-edge noise

Approximations are derived for the three-dimensional, time-harmonic acoustic Green's function whose normal derivative vanishes on the surface of an airfoil of finite thickness and chord for source locations in the neighbourhood of either the leading or trailing edge. The acoustic wavelength is assumed to be large relative to the airfoil thickness, but no restriction is placed on its magnitude relative to . A multiple scattering calculation is performed for high frequencies that involves an expansion in terms of the successive scattering of waves from the leading and trailing edges of the airfoil. The 'principal subseries' of the expansion is summed and shown to provide an excellent approximation for the Green's function when κ 0 ≥ 1, where κ 0 is the acoustic wavenumber. The solution is extended down to κ 0 = 0 by interpolation with the corresponding Green's function for an airfoil of acoustically compact chord. The results extend the single scattering approximation introduced by R. K. Amiet (AIAAJ. 12 1970), and are illustrated by application to the problem of trailing-edge noise generated by nominally steady, low Mach number flow past the airfoil. Experiments and numerical simulations of such flows often include acoustic frequencies that are sufficiently small that the usual assumption of trailing-edge noise theory, that the airfoil is semi-infinite, is not valid.

TRAILING EDGE NOISE AT LOW MACH NUMBERS

A review is made of the diffraction theory of the trailing edge noise generated by a flat-plate airfoil of zero-thickness and non-compact chord, according to which the sound is attributed to the scattering of a “frozen” pattern of turbulence wall pressure swept over the edge in the mean flow. Extension is made to determine the sound produced by very low Mach number flow over the edge of an airfoil of finite thickness. In applications it is desirable to represent the noise in terms of a surface integral over the airfoil involving a Green's function and a metric of the edge flow that can be calculated locally using the equations of motion of an incompressible fluid. It is argued that the appropriate metric for a rigid airfoil is the incompressible “upwash” velocity (determined by the Biot–Savart induction formula applied to the boundary layer vorticity outside the viscous sublayer), and not the surface pressure. Formulae for calculating the noise are given when the airfoil thickness is acoustically compact, and for both three- and two-dimensional edge flows.The theory is illustrated by a detailed discussion of a two-dimensional vortex flow over an airfoil with a rounded trailing edge. The problem is simple enough to be treated analytically, yet is also suitable for validating computational edge noise schemes.

A criterion for leading-edge separation

The maximum angle of attack for unseparated flow over an airfoil of chord c with finite nose radius of curvature r is shown to be 0.818(r/c)½

Effect of temperature on surface noise

An experimental work is discussed whose objective was to obtain data that show the effect of temperature and temperature fluctuations on surface noise. This was accomplished experimentally by immersing a small chord airfoil in the turbulent airstream of a hot jet. The theory and experiment reported by Olsen (1976) provided a guide for designing and validating the hot jet experiment and for interpreting the data. It is shown that increased temperature causes a small decrease in the sound levels; at the same time it causes a shift in the spectra that is smaller but similar to the shift observed with subsonic hot jet noise.

Space-time correlations of airfoil noise

Space-time correlation measurements of surface pressure fluctuations on airfoils, immersed in a cold air jet, with their farfield sound were made over the exit velocity range of 68–190 m/sec (Mach 0.2–0.6). Measurements of surface pressure fluctuations were taken both at the one-third chord and the trailing edge of a 1-in. chord airfoil. The position of the farfield microphone was varied throughout the hemispherical region downstream of the jet exit. Analysis of the cross-correlation and delay time measurements on a 1-in. chord airfoil indicate that the dominant source region for the radiated sound is near the trailing edge. However, for the probe microphone mounted 1 in. from the leading edge of a second, 3-in. chord airfoil, two correlation peaks were recorded. The first corresponds to direct radiation from the probe location, and the second is due to acoustic transmission upstream from the trailing edge. The normalized cross-correlation coefficients were in a ratio of 1:5, respectively. This indicates that the sound from midchord is down 14 dB from the trailing edge sound; theory would indicate even less midchord sound. [Work partially supported by NASA and the TRW Foundation.]

The influence of vortex shedding on the generation of sound by convected turbulence

This paper discusses the theory of the generation of sound which occurs when a frozen turbulent eddy is convected in a mean flow past an airfoil or a semi-infinite plate, with and without the application of a Kutta condition and with and without the presence of a mean vortex sheet in the wake. A sequence of two-dimensional mathematical problems involving a prototype eddy in the form of a line vortex is examined, it being argued that this constitutes the simplest realistic model. Important effects of convection are deduced which hitherto have not been revealed by analyses which assume quadrupole sources to be at rest relative to the plate or airfoil. It is concluded that, to the order of approximation to which the sound from convected turbulence near a scattering body is usually estimated, the imposition of a Kutta condition at the trailing edge leads to a complete cancellation of the sound generated when frozen turbulence convects past a semi-infinite plate, and to the cancellation of the diffraction field produced by the trailing edge in the case of an airfoil of compact chord.

Wake-related sound generation from isolated airfoils

A model for the prediction of wake-related sound generation by a single airfoil is presented which assumes that the net force fluctuation on an airfoil can be expressed in terms of the net momentum fluctuation in the near wake of the airfoil. The model predicts a forcing function for sound generation which is related to the spectra of the two-point velocity correlations in the turbulent region near the airfoil trailing edge. Cross spectra were obtained from correlations of the longitudinal and transverse components of turbulence in the wake of a 36-in. chord NACA series 63-009 airfoil in the 7×10 ft tunnel at the NASA Ames Research Center using x-probe hot-wire sensors. Additional data were obtained from a 10-in. chord airfoil in a free-jet facility at the University of Washington. Based on results from these tests, a scaling model was developed using the turbulent boundary-layer thickness as the primary length measure. The model was used to predict the sound radiated from a 2-in. chord airfoil for which acoustical data is available with good agreement both in level and spectral shape. The single airfoil result is extended to a rotor geometry and comparisons studied for various aerodynamic parameters. Subject Classification:[43]28.65; [43]50.55.

Tunnel interference reduction on a finite airfoil

HE ventilated wind tunnel has been introduced and refined over the last two decades from the fixed geometry to the variable porosity perforated walls. Both theoretical results1 and experimental data2 have demonstrated that it was difficult to eliminate pitching-moment interference of an aircraft model simultaneously with lift interference when using walls with uniformly distributed porosity. The evidence3 in the experimental development of walls for V/STOL testing indicates that it becomes possible to simultaneously eliminate pitching moment and lift interference in a slotted tunnel with nonuniformly distributed porosity. The interference reduction concept has been demonstrated by the author's previous paper4 in which a mathematical technique is presented to predict the interference on an airfoil represented by a single singularity in a Gaussian-type distribution of porosity. The present paper extends the mathematical technique to the case of a finite chord airfoil to predict the proper porosity distribution to eliminate the interference. The present calculation becomes urgent and necessary for an experimental program developing a new wall. On the other hand, the self-correcting wind tunnel concept5'6 which refers to the adaptive wall, has been demonstrated by a numerical method7 to achieve interference-free flow conditions. This wind tunnel requires the provisions either to adjust the deflections of the wall surfaces, or to adjust the wall porosity and plenum pressure behind the wall to establish unconfined flow conditions. The present result of interference calculation is primarily for the improvement of the current existing nonadaptive wall wind tunnel. Analysis The airfoil is placed at the centerline of a perforated tunnel having walls with nonuniform distribution of porosities. The tunnel interference on a model is divided into two parts in the investigation, i.e. the blockage and lift interferences. Based on the thin airfoil theory, the airfoil thickness and camber may be treated separately. The blockage and lift interferences are induced by the presence of the airfoil thickness and camber in the tunnel, respectively. The formulation for both types of interferences is essentially similar, except for the different mathematical representations of the airfoil thickness and camber. The small perturbation theory for subsonic flow is used in the analysis with a homogeneous boundary conModel in a Tunnel