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Showing papers on "Airfoil published in 1994"


Journal ArticleDOI
TL;DR: In this article, the state-space representation of aerodynamic forces and moments for unsteady aircraft motion is proposed, considering separated flow about an airfoil and flow with vortex breakdown about a slender delta wing gives the base for mathematical modeling using internal variables describing the flow state.
Abstract: Mathematical modeling of unsteady aerodynamic forces and moments plays an important role in aircraft dynamics investigation and stability analysis at high angles of attack. In this article the state-space representation of aerodynamic forces and moments for unsteady aircraft motion is proposed. Consideration of separated flow about an airfoil and flow with vortex breakdown about a slender delta wing gives the base for mathematical modeling using internal variables describing the flow state. Coordinates of separation points or vortex breakdown can be taken, e.g., as internal state-space variables. These variables are governed by some differential equations. Within the framework of the proposed mathematical model it is possible to achieve good agreement with different experimental data obtained in water and wind tunnels. These high angle-of-attack experimental results demonstrate considerable dependence of aerodynamic loads on motion time history.

427 citations


Journal ArticleDOI
TL;DR: In this article, surface pressure distributions and wake profiles were obtained for an NACA 4412 airfoil to determine the lift, drag, and pitching-moment coefficients for various configurations.
Abstract: Experimental measurements of surface pressure distributions and wake profiles were obtained for an NACA 4412 airfoil to determine the lift, drag, and pitching-moment coefficients for various configurations. The addition of a Gurney flap increased the maximum lift coefficient from 1.49 up to 1.96, and decreased the drag near the maximum lift condition. There was, however, a drag increment at low-to-moderate lift coefficients. Additional nose-down pitching moment was also generated by increasing the Gurney flap height. Good correlation was observed between the experiment and Navier-Stokes computations of the airfoil with a Gurney flap. Two deploy able configurations were also tested with the hinge line forward of the trailing edge by one and 1.5 flap heights, respectively. These configurations provided performance comparable to that of the Gurney flap. The application of vortex generators to the baseline airfoil delayed boundary-layer separation and yielded an increase in the maximum lift coefficient of 0.34. In addition, there was a significant drag penalty associated with the vortex generators, which suggests that they should be placed where they will be concealed during cruise. The two devices were also shown to work well in concert.

221 citations


01 Jan 1994
TL;DR: The Comprehensive Analytical Model of Rotorcraft Aerodynamics (CAMRAD) as discussed by the authors was developed to calculate rotor performance, loads, and noise; helicopter vibration and gust response; flight dynamics and handling qualities; and system aeroelastic stability.
Abstract: The Comprehensive Analytical Model of Rotorcraft Aerodynamics, CAMRAD, program is designed to calculate rotor performance, loads, and noise; helicopter vibration and gust response; flight dynamics and handling qualities; and system aeroelastic stability. The analysis is a consistent combination of structural, inertial, and aerodynamic models applicable to a wide range of problems and a wide class of vehicles. The CAMRAD analysis can be applied to articulated, hingeless, gimballed, and teetering rotors with an arbitrary number of blades. The rotor degrees of freedom included are blade/flap bending, rigid pitch and elastic torsion, and optionally gimbal or teeter motion. General two-rotor aircrafts can be modeled. Single main-rotor and tandem helicopter and sideby-side tilting proprotor aircraft configurations can be considered. The case of a rotor or helicopter in a wind tunnel can also be modeled. The aircraft degrees of freedom included are the six rigid body motion, elastic airframe motions, and the rotor/engine speed perturbations. CAMRAD calculates the load and motion of helicopters and airframes in two stages. First the trim solution is obtained; then the flutter, flight dynamics, and/or transient behavior can be calculated. The trim operating conditions considered include level flight, steady climb or descent, and steady turns. The analysis of the rotor includes nonlinear inertial and aerodynamic models, applicable to large blade angles and a high inflow ratio, The rotor aerodynamic model is based on two-dimensional steady airfoil characteristics with corrections for three-dimensional and unsteady flow effects, including a dynamic stall model. In the flutter analysis, the matrices are constructed that describe the linear differential equations of motion, and the equations are analyzed. In the flight dynamics analysis, the stability derivatives are calculated and the matrices are constructed that describe the linear differential equations of motion. These equations are analyzed. In the transient analysis, the rigid body equations of motion are numerically integrated, for a prescribed transient gust or control input. The CAMRAD program product is available by license for a period of ten years to domestic U.S. licensees. The licensed program product includes the CAMRAD source code, command procedures, sample applications, and one set of supporting documentation. Copies of the documentation may be purchased separately at the price indicated below. CAMRAD is written in FORTRAN 77 for the DEC VAX under VMS 4.6 with a recommended core memory of 4.04 megabytes. The DISSPLA package is necessary for graphical output. CAMRAD was developed in 1980.

221 citations


Journal ArticleDOI
Kenneth C. Hall1
TL;DR: In this article, a general technique for constructing reduced order models of unsteady aerodynamic flows about two-dimensional isolated airfoils, cascades of airfoil, and three-dimensional wings is developed.
Abstract: A general technique for constructing reduced order models of unsteady aerodynamic flows about twodimensional isolated airfoils, cascades of airfoils, and three-dimensional wings is developed. The starting point is a time domain computational model of the unsteady small disturbance flow. For illustration purposes, we apply the technique to an unsteady incompressible vortex lattice model. The eigenmodes of the system, which may be thought of as aerodynamic states, are computed and subsequently used to construct computationally efficient, reduced order models of the unsteady flowfield. Only a handful of the most dominant eigenmodes are retained in the reduced order model. The effect of the remaining eigenmodes is included approximately using a static correction technique. An important advantage of the present method is that once the eigenmode information has been computed, reduced order models can be constructed for any number of arbitrary modes of airfoil motion very inexpensively. Numerical examples are presented that demonstrate the accuracy and computational efficiency of the present method. Finally, we show how the reduced order model may be incorporated into an aeroelastic flutter model.

206 citations


Journal ArticleDOI
TL;DR: In this article, boundary-layer separation control on a twodimensional single-flap, three-element, high-lift system at near-flight Reynolds numbers with small surfacemounted vortex generators is evaluated.
Abstract: An experimental investigation has been conducted to evaluate boundary-layer separation control on a twodimensional single-flap, three-element, high-lift system at near-flight Reynolds numbers with small surfacemounted vortex generators. The wind-tunnel testing was carried out in the NASA Langley Low-Turbulence Pressure Tunnel as part of a cooperative program between McDonnell Douglas Aerospace and NASA Langley Research Center to develop code validation data bases and to improve physical understanding of multielement airfoil flows. This article describes results obtained for small (subboundary-layer) vane-type vortex generators mounted on a multielement airfoil in a landing configuration. Measurements include lift, drag, surface pressure, wake profile, and fluctuating surface heat fluxes. The results reveal that vortex generators as small as 0.18% of reference (slat and flap stowed) wing chord ("micro-vortex generators") can effectively reduce boundarylayer separation on the flap for landing configurations. Reduction of flap separation can significantly improve performance of the high-lift system by reducing drag and increasing lift for a given approach angle of attack. At their optimum chordwise placement on the flap, the micro-vortex generators are hidden inside the wing when the flap is retracted, thus extracting no cruise drag penalty.

169 citations


01 Sep 1994
TL;DR: In this article, a comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted Testing the wing in the nonrotating condition isolates the three-dimensional (3D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment.
Abstract: A comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted Testing the wing in the nonrotating condition isolates the three-dimensional (3-D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment The test has generated a very complete, detailed, and accurate body of data These data include static and dynamic pressure distributions, surface flow visualizations, two-dimensional (2-D) airfoil data from the same model and installation, and important supporting blockage and wall pressure distributions This body of data is sufficiently comprehensive and accurate that it can be used for the validation of rotor blade aerodynamic models over a broad range of the important parameters including 3-D dynamic stall This data report presents all the cycle-averaged lift, drag, and pitching moment coefficient data versus angle of attack obtained from the instantaneous pressure data for the 3-D wing and the 2-D airfoil Also presented are examples of the following: cycle-to-cycle variations occurring for incipient or lightly stalled conditions; 3-D surface flow visualizations; supporting blockage and wall pressure distributions; and underlying detailed pressure results

125 citations


Journal ArticleDOI
TL;DR: In this paper, a new approach for combining conceptual and preliminary design techniques for wing optimization is presented for the high-speed civil transport (HSCT) and a wing shape parametrization procedure is developed which allows the linking of planform and airfoil design variables.
Abstract: A new approach for combining conceptual and preliminary design techniques for wing optimization is presented for the high-speed civil transport (HSCT). A wing-shape parametrization procedure is developed which allows the linking of planform and airfoil design variables. Variable-complexity design strategies are used to combine conceptual and preliminary-design approaches, both to preserve interdisciplinary design influences and to reduce computational expense. In the study, conceptual-design-level algebraic equations are used to estimate aircraft weight, supersonic wave drag, friction drag, and drag due to lift. The drag due to lift and wave drag are also evaluated using more detailed, preliminary-design-level techniques. The methodology is applied to the minimization of the gross weight of an HSCT that flies at Mach 3 with a range of 6500 mi.

123 citations


Journal ArticleDOI
TL;DR: The ability of one-and two-equation turbulence models to predict unsteady separated flows over airfoils is evaluated in this paper, where an implicit, factorized, upwindbiased numerical scheme is used for the integration of the compressible, Reynolds-averaged Navier-Stokes equations.
Abstract: The ability of one- and two-equation turbulence models to predict unsteady separated flows over airfoils is evaluated. An implicit, factorized, upwind-biased numerical scheme is used for the integration of the compressible, Reynolds-averaged Navier-Stokes equations. The turbulent eddy viscosity is obtained from the computed mean flowfield by integration of the turbulent field equations. One- and two-equation turbulence models are first tested for a separated airfoil flow at fixed angle of incidence. The same models are then applied to compute the unsteady flowfields about airfoils undergoing oscillatory motion at low subsonic Mach numbers. Experimental cases where the flow has been tripped at the leading-edge and where natural transition was allowed to occur naturally are considered. The more recently developed turbulence models capture the physics of unsteady separated flow significantly better than the standard kappa-epsilon and kappa-omega models. However, certain differences in the hysteresis effects are observed. For an untripped high-Reynolds-number flow, it was found necessary to take into account the leading-edge transitional flow region to capture the correct physical mechanism that leads to dynamic stall.

119 citations


Journal ArticleDOI
TL;DR: In this paper, the authors describe massively parallel finite element computations of unsteady incompressible flows involving fluid-body interactions, based on the Deforming-Spatial-Domain/Stabilized-Space-Time (DSD/SST) finite element formulation.

115 citations


Journal ArticleDOI
TL;DR: In this paper, a practical method for computing the unsteady lift on an airfoil due to arbitrary motion of a trailing-edge flap was described, and the result for the incompressible case was obtained in state-space form by means of Duhamel superposition and employing an improved exponential approximation to Wagner's indicia! lift function.
Abstract: A practical method is described for computing the unsteady lift on an airfoil due to arbitrary motion of a trailing-edge flap. The result for the incompressible case is obtained in state-space form by means of Duhamel superposition and employing an improved exponential approximation to Wagner's indicia! lift function. For subsonic compressible flow, the indicial lift at small values of time due to impulsive trailing-edge flap deflection is obtained from linear theory in conjunction with the aerodynamic reciprocal theorems. These results are used with experimental results for the oscillating case to obtain complete exponential approximations for the indicial response due to impulsive flap deflection. The final result for the unsteady lift due to an arbitrary flap motion in subsonic flow is obtained in state-space form. Numerical results and comparisons with experimental data are shown.

112 citations


Journal ArticleDOI
TL;DR: In this article, a two-dimensional airfoil with a free-play nonlinearity in pitch subject to incompressibl e flow was analyzed and the aerodynamic forces were evaluated using Wagner's function and the resulting equations integrated numerically to give time histories of the air-foil motion.
Abstract: A two-dimensiona l airfoil with a free-play nonlinearity in pitch subject to incompressibl e flow has been analyzed. The aerodynamic forces on the airfoil were evaluated using Wagner's function and the resulting equations integrated numerically to give time histories of the airfoil motion. Regions of limit cycle oscillation are detected for velocities well below the linear flutter boundary, and the existence of these regions is strongly dependent on the initial conditions and properties of the airfoil. Furthermore, for small structural preloads, narrow regions of chaotic motion are obtained, as suggested by power spectral densities, phase-plane plots, and Poincare sections of the airfoil time histories. The existence of this chaotic motion is strongly dependent on a number of airfoil parameters, including, mass, frequency ratio, structural damping, and preload.

01 Jan 1994
TL;DR: In this article, a two-dimensional airfoil with a free-play nonlinearity in pitch subject to incompressibl e flow was analyzed and the aerodynamic forces were evaluated using Wagner's function and the resulting equations integrated numerically to give time histories of the air-foil motion.
Abstract: A two-dimensiona l airfoil with a free-play nonlinearity in pitch subject to incompressibl e flow has been analyzed. The aerodynamic forces on the airfoil were evaluated using Wagner's function and the resulting equations integrated numerically to give time histories of the airfoil motion. Regions of limit cycle oscillation are detected for velocities well below the linear flutter boundary, and the existence of these regions is strongly dependent on the initial conditions and properties of the airfoil. Furthermore, for small structural preloads, narrow regions of chaotic motion are obtained, as suggested by power spectral densities, phase-plane plots, and Poincare sections of the airfoil time histories. The existence of this chaotic motion is strongly dependent on a number of airfoil parameters, including, mass, frequency ratio, structural damping, and preload.

Journal ArticleDOI
TL;DR: In this paper, the lift hysteresis loops are estimated from the shed vorticity flux and the analytical foundation of the method and the various approximations are discussed.
Abstract: The flow field of an airfoil oscillated periodically over a reduced frequency range, 0 ≤ k ≤ 1.6, is studied experimentally at chord Reynolds numbers of Rc = 22000 and 44000. For most of the data, the NACA0012 airfoil is pitched sinusoidally about one quarter chord between angles of attack α of 5° and 25°. The cyclic variation of the near wake flow field is documented through flow visualization and phase-averaged vorticity measurements. In addition to the familiar dynamic stall vortex (DSV), an intense vortex of opposite sign is observed to originate from the trailing edge just when the DSV is shed. The two together take the shape of the cross-section of a large ‘mushroom’ while being convected away from the airfoil. The phase delay in the shedding of the DSV with increasing k, as observed by previous researchers, is documented for the full range of k. It is observed that the sum of the absolute values of all vorticity convected into the wake over a cycle is nearly constant and is independent of the reduced frequency and amplitude of oscillation but dependent on the mean α. The time varying component of the lift is estimated in a novel way from the shed vorticity flux. The analytical foundation of the method and the various approximations are discussed. The estimated lift hysteresis loops are found to be in reasonable agreement with available data from the literature as well as with limited force balance measurements. Comparison of the lift hysteresis loops with the corresponding vorticity fields clearly shows that major features of the lift variation are directly linked to the evolution of the large-scale vortical structures and the phase delay phenomenon.

Journal ArticleDOI
TL;DR: In this paper, the initial stages of two-dimensional al unsteady leading-edge boundary-layer separation of laminar subsonic flow over a pitching NACA-0012 airfoil have been studied numerically at Reynolds number (based on air-foil chord length) Rec = 104, Mach number Mx = 0.2, and non-dimensional pitch rate H£ =0.2.
Abstract: The initial stages of two-dimension al unsteady leading-edge boundary-layer separation of laminar subsonic flow over a pitching NACA-0012 airfoil have been studied numerically at Reynolds number (based on airfoil chord length) Rec = 104, Mach number Mx = 0.2, and nondimensional pitch rate H£ = 0.2. Computations have been performed using two separate algorithms for the compressible laminar Navier-Stokes equations. The first method, denoted the structured grid algorithm, utilizes a structured, boundary-fitted C grid and employs the implicit approximate-factorization algorithm of Beam and Warming. The second method, denoted the unstructured grid algorithm, utilizes an unstructured grid of triangles and employs the flux-difference splitting method of Roe and a discrete representation of Gauss' theorem for the in viscid and viscous terms, respectively. Both algorithms are second-order accurate in space and time and have been extensively validated through comparison with analytical and previous numerical results for a variety of problems. The results show the emergence of a primary clockwise-rotating recirculating region near the leading edge which can be traced to a pair of critical points (a center and a saddle) that appear within the flowfield, followed by a secondary counterclockwise-rotating recirculating region and a tertiary clockwise-rotating recirculating region. The primary and secondary recirculating regions interact with each other to give rise to the unsteady separation ("breakaway") of the boundary layer.

Patent
28 Jun 1994
TL;DR: In this paper, an inner and/or outer cylindrical wall limiting the working fluid flow path of an axial compressor radially has an ondulating contour, where at the intersection with the leading edge of an airfoil the wall shows a convex contour followed by a concave contour in the region of the airfoils maximum thickness.
Abstract: An inner and/or outer cylindrical wall limiting the working fluid flow path of an axial compressor radially has an ondulating contour. At the intersection with the leading edge of an airfoil the wall shows a convex contour (54) followed by a concave contour (58) in the region of the airfoils maximum thickness while at the intersection with the trailing edge of the airfoil the contour (56) is convex again. The airfoil can either be a rotor blade or a stator vane.

Journal ArticleDOI
TL;DR: Theoretical and practical aspects of conducting three-dimensional wake measurements in large wind tunnels are reviewed with emphasis on applications in low-speed aerodynamics in this paper, where the authors demonstrate the value of this measurement technique using data from wake measurements conducted on a variety of low speed configurations including the complex high-lift system of a transport aircraft.
Abstract: Theoretical and practical aspects of conducting three-dimensional wake measurements in large wind tunnels are reviewed with emphasis on applications in low-speed aerodynamics. Such quantitative wake surveys furnish separate values for the components of drag such as profile drag and induced drag, but also measure lift without the use of a balance. In addition to global data, details of the wake flowfield as well as spanwise distributions of lift and drag are obtained. This article demonstrates the value of this measurement technique using data from wake measurements conducted on a variety of low-speed configurations including the complex high-lift system of a transport aircraft.

Journal ArticleDOI
TL;DR: In this paper, the effects of a periodic free-stream velocity on the aerodynamic properties of an airfoil in incompressible flow are examined, and a new general aerodynamic theory for a combination of harmonic pitching, plunging and fore-aft motion is presented.
Abstract: The effects of a periodic free-stream velocity on the unsteady aerodynamics of an airfoil in incompressible flow are examined. Existing theories are reviewed, and their simplifications and limitations are properly identified. A new general aerodynamic theory for an airfoil undergoing a combination of harmonic pitching, plunging and fore-aft motion is presented. An extension to arbitrary free-stream velocity variations and arbitrary airfoil motion is also given. The theoretical results are validated against numerical predictions made by a modern Euler code.

Patent
15 Dec 1994
TL;DR: In this paper, a series of vortex generators are provided each of which comprises a cavity in the component over which the supersonic air is flowing that is configured to generate a spiral vortex which attenuates flow separation and weight drag resulting from the SUpersonic airflow.
Abstract: A vortex generator for attenuating flow separation which occur during supersonic flow of air over structure such as an aircraft airfoil, its fuselage, surfaces forming a part of a jet propulsion unit, turbine or compressor blades, or similar surfaces subjected to supersonic airflow. A series of vortex generators are provided each of which comprises a cavity in the component over which the supersonic air is flowing that is configured to generate a spiral vortex which attenuates flow separation and weight drag resulting from the supersonic airflow. Each cavity is of generally triangular configuration defined by two side walls which diverge in a direction away from the apex of the triangular cavity, and a flat bottom wall joined to the side walls. In an alternate embodiment, means is provided for selectively shifting the bottom wall from a retracted inner position, to an outer location essentially flush with the surface over which the supersonic airflow is occurring.

01 Dec 1994
TL;DR: In this article, an improved model of stalled flow on rotating blades in a time averaged sense, in order to predict rotor performance in the stall regime, was established through the following steps, all of which are discussed in this report: (a) analysis of the boundary layer (b.l.) equations for a reference system rotating with the blades for attached and for stalled flow; (b) formulation of a quasi 3D system of b.l. equations that include the leading terms due to rotation, both for the stalled case as for the attached case; this set of equations
Abstract: Aim of the investigations was to establish improved modelling of stalled flow on rotating blades in a time averaged sense, in order to predict rotor performance in the stall regime. This is achieved through the following steps, all of which are discussed in this report: (a) analysis of the boundary layer (b.l.) equations for a reference system rotating with the blades for attached and for stalled flow; (b) formulation of a quasi 3D system of b.l. equations that include the leading terms due to rotation, both for the stalled case as for the attached case; this set of equations permits strip wise calculations; (c) extension of the NLR (2D) ULTRAN-V viscous inviscid interaction code to accommodate radial flow and the 3D terms in the chord wise momentum equation; (d) implementation of the extended boundary layer equations in ULTRAN-V; (e) analysis of the FFA measurement in the CARDC tunnel in order to obtain c[sub 1]-[alpha] data at different sections on the rotating blade; (f) calculations with the extended code for the FFA configuration and comparison between calculated and measured results; (g) analysis of the differences and determination of a calibration constant; (h) determination of a simple correction formula to obtain rotating c[sub 1]-[alpha] data from measured 2D data, based on curve fitting of calculated results; (i) comparison of data obtained with the method developed with measurements executed on the DUT test rotor. Finally a number of power curves calculated with the synthesized 3D c[sub 1]-[alpha] data are compared with measurements. Results produced with this input are much closer to measurements than calculated power curves with 2D airfoil date. Parts of the results discussed in this report were published in conference proceedings, during the course of the investigations. In the present report comprehensive view of the modelling efforts of the entire project is given. 24 figs., 4 tabs., 2 appendices, 24 refs.

Proceedings ArticleDOI
10 Jan 1994
TL;DR: In this article, four different turbulence models are used to compute the flow over a three-element airfoil configuration, including the Baldwin-Barth model, the Spalart-Allmaras model, a two-equation k-omega model, and a new Durbin-Mansour model.
Abstract: Four different turbulence models are used to compute the flow over a three-element airfoil configuration. These models are the one-equation Baldwin-Barth model, the one-equation Spalart-Allmaras model, a two-equation k-omega model, and a new one-equation Durbin-Mansour model. The flow is computed using the INS2D two-dimensional incompressible Navier-Stokes solver. An overset Chimera grid approach is utilized. Grid resolution tests are presented, and manual solution-adaptation of the grid was performed. The performance of each of the models is evaluated for test cases involving different angles-of-attack, Reynolds numbers, and flap riggings. The resulting surface pressure coefficients, skin friction, velocity profiles, and lift, drag, and moment coefficients are compared with experimental data. The models produce very similar results in most cases. Excellent agreement between computational and experimental surface pressures was observed, but only moderately good agreement was seen in the velocity profile data. In general, the difference between the predictions of the different models was less than the difference between the computational and experimental data.

Journal ArticleDOI
TL;DR: In this paper, an experimental/analytical research program was undertaken to develop advanced versions of circulation control wing (CCW) blown high-lift airfoils, and to address specific issues related to their application to subsonic transport aircraft.
Abstract: An experimental/ analytical research program was undertaken to develop advanced versions of circulation control wing (CCW) blown high-lift airfoils, and to address specific issues related to their application to subsonic transport aircraft. The primary goal was to determine the feasibility and potential of these pneumatic airfoils to increase high-lift system performance in the terminal area while reducing system complexity. A four-phase program was completed, including 1) experimental development and evaluation of advanced CCW high-lift configurations, 2) development of effective pneumatic leading-edge devices, 3) computational evaluation of CCW airfoil designs plus high-lift and cruise capabilities, and 4) investigation of the terminal-area performance of transport aircraft employing these airfoils. The first three phases of this program are described in Part I of this article. Applications to the high-lift and control systems of advanced subsonic transport aircraft and resulting performance are discussed in the continuation of this article, Part II. Experimental lift coefficient values approaching 8.0 at zero incidence and low blowing rates were demonstrated by two-dimensional CCW configurations that promised minimal degradation of the airfoil's performance during cruise. These results and experimental/CFD methods will be presented in greater detail in the following discussions.

01 Aug 1994
TL;DR: In this paper, the authors describe the implementation of optimization techniques based on control theory for airfoil and wing design, and present results for both two and three dimensional cases, including the optimization of a swept wing.
Abstract: These lectures describe the implementation of optimization techniques based on control theory for airfoil and wing design. In previous studies it was shown that control theory could be used to devise an effective optimization procedure for two-dimensional profiles in which the shape is determined by a conformal transformation from a unit circle, and the control is the mapping function. Recently the method has been implemented in an alternative formulation which does not depend on conformal mapping, so that it can more easily be extended to treat general configurations. The method has also been extended to treat the Euler equations, and results are presented for both two and three dimensional cases, including the optimization of a swept wing.

Journal ArticleDOI
TL;DR: In this paper, the dynamic-stall vortex (DSV) was suppressed by removing an appropriate amount of the reverse-flowing fluid to prevent its accumulation in the near-leading edge region, thereby preventing lift up of the shear layer.
Abstract: Experiments to control the dynamic-stall vortex (DSV) over the suction surface of a two-dimensional NACA 0012 airfoil, undergoing a hold-pitch-hold motion, are described. Measurements were performed over a range of Reynolds number (3.0×10 4 ≤Re c ≤1.18×10 5 ) and pitch rate (0.072≤α + ≤0.31), using leading-edge suction duing a prescribed period of the airfoil motion. This strategy to manage the DSV, using controlled leading-edge suction, was developed from a study of the mechanisms responsible for the evolution of the vortex. The results indicate that formation of the DSV can be suppressed by removing an appropriate amount of the reverse-flowing fluid to prevent its accumulation in the near-leading-edge region, thereby preventing lift up of the shear layer

Patent
22 Aug 1994
TL;DR: In this paper, a steam turbine blade having an airfoil portion and root portion by which the blade is affixed to a rotor is configured to minimize energy loss through the row of blades.
Abstract: A steam turbine blade having an airfoil portion and root portion by which the blade is affixed to a rotor. The geometry of the blade airfoil is configured to minimize energy loss through the row of blades and reduce the weight of the airfoil. The airfoil has a leading edge and a trailing edge defining a chord therebetween. The chord is reduced linearly from the base of the airfoil to 50% of the airfoil height. However, the chord remains essentially constant from 50% of the airfoil height to the airfoil tip. The root is fir tree shaped and has four sets of tangs and grooves that are configured to minimize the stresses in the root.

Patent
01 Nov 1994
TL;DR: In this article, the optimal stiffness parameters for airfoils with small geometric features, such as the crossover holes, are provided to improve the overall cooling scheme without jeopardizing manufacturability of airfoILS.
Abstract: An airfoil (20) for a gas turbine engine (10) includes cooling passages (40), (50) extending radially within the airfoil to circulate cooling air therethrough. Pluralities of small crossover holes (48), (66), (72) are formed within the walls (50), (68), (74), respectively, to allow cooling air to flow between the cooling passages. Optimum stiffness parameters are provided to improve producability of the airfoils with small geometric features, such as the crossover holes, as well as to improve the overall cooling scheme without jeopardizing manufacturability of airfoils.

Patent
07 Dec 1994
TL;DR: In this paper, methods for improving the erosion resistance of composite airfoils are disclosed as well as the resultant structures, where wire mesh materials are coated with an erosion-resistant coating, formed to the shape of the airfoil leading edge, and molded into the leading edge during air-foil fabrication.
Abstract: Methods for improving the erosion resistance of composite airfoils are disclosed as are the resultant structures. Wire mesh materials are coated with an erosion-resistant coating, formed to the shape of the airfoil leading edge, and molded into the leading edge during airfoil fabrication.

Patent
01 Jul 1994
TL;DR: In this paper, a coordinate measuring machine measures external coordinates along an airfoil and sends the coordinates to a computer, where the computer sorts and orders the coordinates into a plurality of probe center points.
Abstract: The present invention discloses a system and method for determining airfoil characteristics from coordinate measuring machine probe center data. A coordinate measuring machine measures external coordinates along an airfoil and sends the coordinates to a computer. The computer sorts and orders the coordinates into a plurality of probe center points along a plurality of sections along the airfoil. The probe center points are triangulated to detect and eliminate errant points. Then the probe center points are correlated to a nominal part of the airfoil stored in the memory of the computer until a specified plane of interest is obtained. A discrete inset operation is performed on the correlated data for the plane of interest until a maximum thickness is obtained. The inset points are then joined to form an airfoil meanline, which is used to determine characteristics such as leading edge thickness, trailing edge thickness, and chord length.

Book ChapterDOI
20 Jun 1994
TL;DR: In this article, the ASU Unsteady Wind Tunnel on a 45° swept airfoil was used to measure roughness-induced stationary crossflow in the presence of roughness elements.
Abstract: Stability experiments are conducted in the ASU Unsteady Wind Tunnel on a 45° swept airfoil. The surface of the is polished to 0.25 μm rms. Under these conditions, natural stationary crossflow vortices are not measurable. This state is used to measure roughness-induced stationary crossflow. Spanwise arrays of 70–150 μm roughness elements are introduced near the attachment line. Detailed hot-wire measurements are taken to document the growth of these vortices. The data clearly show that linear stability theory does not accurately predict the growth rates of stationary crossflow waves under these conditions.

Patent
17 Oct 1994
TL;DR: In this article, a control system for aircraft airfoils, which is an improvement over existing aileron, flap, spoiler and deicing technologies, in providing increased roll control and aerodynamic lifting and braking functions, with greatly reduced drag increased airspeed and precise control performance at all airspeeds, due to clean uninterrupted airfoil surfaces and directional conformance of wing to the intended flight path.
Abstract: A control system for aircraft airfoils, which is an improvement over existing aileron, flap, spoiler and deicing technologies, in providing increased roll control and aerodynamic lifting and braking functions; with greatly reduced drag increased airspeed and precise control performance at all airspeeds, due to clean uninterrupted airfoil surfaces and directional conformance of wing to the intended flight path. This is accomplished by use of a torque tube mounted internally in the aeroelastic airfoil structure, and firmly attached to the airfoil tip structure. In operation the inboard end of the torque tube when rotated differentially on its pivot axis, imposes a helicoidal twist on the aeroelastic airfoil structure, with maximum angle of incidence at the outboard wing tip, providing near perfect lateral roll control or cooperating to provide increased lift and braking or maneuverability, also the foregoing operations provide automatic deicing. The torque tube can be operated by conventional control systems, e.g. cable/pulley, electric/hydraulic servo etc.

Journal ArticleDOI
TL;DR: In this paper, the effects of blade row interaction on the aerodynamics of a transonic turbine stage were investigated using a two-dimensional unsteady Navier-Stokes code based on an explicit Runge-Kutta algorithm and an overlapping O-H grid system.
Abstract: Part I of this article presents results of a computational investigation of the effects of blade row interaction on the aerodynamics of a transonic turbine stage. The predictions are obtained using a two-dimensional unsteady Navier-Stokes code based on an explicit Runge-Kutta algorithm and an overlapping O-H grid system. This code simulates the flow in time-accurate fashion using nonreflective stage inflow and outflow boundary conditions and phase-lagging procedures for modeling arbitrary airfoil counts in the vane and blade rows. The O-H grid provides high spatial resolution of the high gradient regions near the airfoil surfaces and allows for arbitrary placement of stage inflow and outflow boundaries. Unsteady and time-averaged airfoil surface pressure predictions are compared with those from an older version of the code based on the explicit hopscotch algorithm and an O-grid system, and experimental data obtained in a short-duration shock tunnel facility.