Effects Of Camber Research Articles

Interactive Multimode Blade Impact Analysis

This paper describes the theoretical methodology used in developing an analysis for the response of turbine engine fan blades subjected to soft body (bird) impacts and the computer program that was developed using this methodology as its basis. This computer program is an outgrowth of two programs that were previously developed for the purpose of studying problems of a similar nature (a three-mode beam impact analysis and a multimode beam impact analysis). The present program utilizes an improved missile model that is interactively coupled with blade motion which is more consistent with actual observations. It takes into account local deformation at the impact area, blade camber effects, and the spreading of the impacted missile mass on the blade surface. In addition, it accommodates plate-type mode shapes. The analysis capability in this computer program represents a significant improvement in the development of the methodology for evaluating potential fan blade materials and designs with regard to foreign object impact resistance.

Effect of initial camber on natural nonlinear vibrations of cylindrical shells

Effect of initial camber on natural nonlinear vibrations of cylindrical shells

Contributions to the virtual mass theory of hydrodynamic planning

Virtual mass solutions (Payne, 1974, 1980a) for the normal force on a flat planing plate gave good agreement with experiments for wetted length to beam ratios ( l ̂ = l/b) in excess of unity. In an article by Payne (1981a), the region below l ̂ = 1 was faired into the known two-dimensional solution (infinite aspect ratio) by means of a bridging function. A subsequent study (Payne, 1980c) of the virtual mass of a rectangular plate led to another bridging function which seemed to be more fundamental, and might give better agreement between theory and experiment below l ̂ = 1 than was the case in the work by Payne (1981a). In the present note this is shown to be so. And although the lift curve slope of a wing and a planing plate (when the cavity lift contribution is neglected) are identical for slender planforms ( l ̂ > 1) and in the two dimensional flow ( l ̂ = 0) the plate is found to have a higher lift curve slope in between. There is no relationship between the pitching moment of a planing plate and an aerofoil. This subsequent work has also resulted in a more elegant version of the basic equations, and the new version is presented for both normal force and moment. For slender ( l ̂ = 1) flat plates the non-dimensional center of pressure ( X ̂ G) of the “virtual mass” force is close to the stagnation line and moves aft with increasing τ and l ̂ . It is rarely further aft than 10% of the wetted length. The theoretical prediction is compared with the measurements of Sottorf and found to be within the scatter of the data. In contrast, X ̂ G = 2 3 for a chines dry prismatic hull, and 1 4 for a wing or hydrofoil. The theory enables the effect of camber and rocker to be studied. Using simplified approximations to the general theory it is shown that camber reduces pressure drag (by as much as half) and the bow-up pitching moment; the latter by so much that a cambered planing surface can become unstable. (This was already known theoretically for two-dimensional flow, from the work of Wagner (1932a,b), Cumberbatch (1958), Wu and Whitney (1972), Doctors (1974), Wellicome and Jahangeer (1978). It was known for slender planing surfaces from the work of Maruo (1967) and Tulin (1957). It was also known experimentally for finite aspect ratios from the work of Sottorf (1932, 1934), Clements (1964) and Moore (1967). Rocker, on the other hand, gives a large bow-up couple, and results in increased pressure drag. Presumably this large bow-up moment is why rocker is almost invariably employed with planing sailing dinghies and surfboards, despite the resistance penalty.

Discussion: “Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies” (George, A. R., 1981, ASME J. Fluids Eng., 103, pp. 631–637)

Discussion: “Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies” (George, A. R., 1981, ASME J. Fluids Eng., 103, pp. 631–637)

Closure to “Discussions of ‘Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies’” (1981, ASME J. Fluids Eng., 103, pp. 637–638)

Closure to “Discussions of ‘Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies’” (1981, ASME J. Fluids Eng., 103, pp. 637–638)

Discussion: “Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies” (George, A. R., 1981, ASME J. Fluids Eng., 103, pp. 631–637)

Discussion: “Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies” (George, A. R., 1981, ASME J. Fluids Eng., 103, pp. 631–637)

Aerodynamic Effects of Shape, Camber, Pitch, and Ground Proximity on Idealized Ground-Vehicle Bodies

Results are presented from an experimental study of the lift, drag, pitching moment, and flow field of a series of rounded edge simple bluff bodies of various cambers and tapers. The bodies were proportioned to be similar to those of idealized ground vehicles such as automobiles, vans, and trucks. The models were tested with and without simulated wheels, underbody roughness, and proximity to a stationary and moving ground plane. The pitch angle was varied at zero yaw angle. The force and moment coefficients and flow visualization studies indicated the existence and importance of flow regimes characterized by a pair of trailing vortices on the leeward side of the body similar to those found over an inclined body of revolution and over slender delta wings. These flows can suppress bubble-type separation. The effects of a rough underbody are generally detrimental although less so if the rough surface is on the windward side. A moving ground plane was found to give significantly different lift and drag for small ground clearances characteristic of actual road vehicles.

The Time-Variant Aerodynamic Response of a Stator Row Including the Effects of Airfoil Camber

An experimental investigation was conducted to quantitatively determine the validity and applicability of state-of-the-art transverse gust cascade analyses. This was accomplished by obtaining fundamental time-variant forced response data at realistic values of key parameters in a large-scale, low-speed, single-stage research compressor. The forcing function, the velocity defect created by the rotor blade wakes, was measured with a crossed hot-wire probe. The resulting time-variant aerodynamic response was measured by means of flush mounted high response pressure transducers on both flat plate and cambered airfoil stator vane rows over a wide range of incidence angles. These dynamic data were then analyzed to determine the chordwise variation of the unsteady pressure difference in terms of a dimensionless dynamic pressure coefficient and an aerodynamic phase lag referenced to a transverse gust at the leading edge of the vanes. These dimensionless pressure difference data were all correlated with predictions obtained from a state-of-the-art compressible transverse gust, flat plate cascade analysis. Correlation of the classical flat plate unsteady data with the predictions permitted the range of validity of the analysis to be assessed in terms of incidence angle. Correlation of the cambered vane unsteady data with those for the flat plate and with the predictions allowed the effects of airfoil camber as well as the applicability of the flat plate prediction to realistic cambered airfoils to be quantitatively determined.

Closure to “Discussion of ‘The Time-Variant Aerodynamic Response of a Stator Row Including the Effects of Airfoil Camber’” (1980, ASME J. Eng. Power, 102, p. 342)

Closure to “Discussion of ‘The Time-Variant Aerodynamic Response of a Stator Row Including the Effects of Airfoil Camber’” (1980, ASME J. Eng. Power, 102, p. 342)

Discussion: “The Time-Variant Aerodynamic Response of a Stator Row Including the Effects of Airfoil Camber” (Fleeter, S., Bennett, W. A., and Jay, R. L., 1980, ASME J. Eng. Power, 102, pp. 334–342)

Discussion: “The Time-Variant Aerodynamic Response of a Stator Row Including the Effects of Airfoil Camber” (Fleeter, S., Bennett, W. A., and Jay, R. L., 1980, ASME J. Eng. Power, 102, pp. 334–342)

Time-Variant Aerodynamics of Oscillating Airfoil Surfaces in a Supersonic Flowfield

The results of an experimental study of the time-variant aerodynamics associated with harmonically oscillating single airfoil surfaces in a supersonic flowfield are presented. Six single airfoil configurations were investigated: a flat plate, wedge, flat plate-convex corner, wedge-convex corner, MCA suction surface, and MCA pressure surface. The basic time-variant aerodynamic data obtained for each of these configurations included the chordwise distribution of the unsteady pressure magnitude (expressed as an unsteady pressure coefficient) and its phase lag as referenced to the airfoil motion, with Mach number as parameter. All of the data obtained are correlated with a low reduced frequency prediction which includes the effects of airfoil camber and thickness distribution.

The Development of High Lift, Single-Component Airfoil Sections

SummaryThis paper investigates the possibility of increasing the maximum lift of a single element wing section by the introduction of camber near the trailing edge. The optimum upper surface velocity distribution consists of a constant high velocity “rooftop” over the forward part of the wing section followed by transition and an optimised turbulent boundary layer recovery region to the trailing edge. The height of the rooftop, which essentially determines the maximum lift coefficient, is limited by the amount of recovery attainable without separation of the turbulent boundary layer. An increase in the upper surface velocity at the trailing edge would allow the whole top surface velocity distribution to be raised resulting in a considerable increase in maximum lift. The introduction of camber over the rear part of the wing section should produce the required increase in trailing edge velocity on the upper surface. The practical possibilities of using this technique for increased maximum lift were examined by designing a high lift wing section and testing it in the University of Alberta Low Turbulence Wind Tunnel. The results showed that the new wing section behaved qualitatively as expected. A difference of about 0.5U∞ between upper and lower surface velocities at the trailing edge was achieved, even though the thick boundary layers near the trailing edge tend to reduce the effect of camber. A maximum lift coefficient of 2.64 was measured at a Reynolds number of 1.0 × 106.

Subsonic Vortex-Flow Design Study for Slender Wings

A theoretical study describing the effects of spanwise camber on the lift dependent drag of slender delta wings having leading-edge vortex flow is presented. The earlier work by Barsby, using conical flow, indicated that drag levels similar to those in attached flow could be obtained. This is re-examined and then extended to the more practical case of nonconical flow by application of the vortex-lattice method coupled with the suction analogy and the recently developed Boeing free-vortex-sheet method. Lastly, a design code is introduced which employs the suction analogy in an attempt to define optimum camber surfaces for minimum lift dependent drag for vortex flow conditions.

The effect of camber on the pressure field of a super-cavitating propeller

An expression is obtained for the cross-sectional area of a super-cavity on a propeller blade section with camber. With help of this expression the effect of camber on the pressure field induced by the super-cavitating part of a propeller can be inve

Spanwise cambered delta wing with leading-edge separation

Smith's method of calculating flow with leading separation on plane delta wings in subsonic flow is extended to the case of spanwise cambered delta wings, through the use of a system of conformal transformations. A numerical procedure to develop the required transformation is described. Using the transformations, a study of the effect of spanwise camber on the flow characteristics such as vortex core location and strength, sheet shape, pressure distribution, and normal force is made on a conically cambered delta wing. It is found that, at a given incidence measured from the respective zero lift angle, the core vortex on the cambered wing lies more outboard towards the leading edge and closer to the wing surface than that on the plane wing. Core and sheet vortex strengths of the cambered wing are less than those of plane wings. Significant changes in sheet shape and the location and magnitude of suction peaks are observed as a function of spanwise camber. However, with increase in incidence, the camber wing characteristics approach those of the plane wing, indicating the diminishing influence of spanwise camber on the flow details at large incidence. Since the theoretical flow model is based on the assumption of conical flow in the separated flow field and slender wing theory for attached flow, the method is applicable in the forward region near the apex of low aspect ratio wings.

Noise generated by low pressure axial flow fans, II: Effects of number of blades, chord length and camber of blade

In a previous paper [1], a practical formula which correlates the sound pressure level of turbulent noise originating from an axial flow fan to the design parameters of the impeller was derived analytically by introducing a simple flow model. The validity of this formula is examined in respect to the effects of some of these parameters, such as number of blades and chord length, on the noise level. Experiments are also described on the effect of blade camber, to examine the applicability of the formula adopted for the evaluation of the wake width of the blade in the calculation of theoretical values, because the wake width plays a very important role in estimating the sound power. The agreement between the experimental and theoretical results is good unless separation occurs on the blade surface. Finally, it is shown that the measured sound pressure level data for most of the test impellers fall on a single curve when a dimensionless parameter, named the DT ratio (which represents the ratio of the wake width to the pitch of the blade row), is used as the abscissa.

Supercirculation Effects Induced by Vectoring a Partial-Span Rectangular Jet

Thrust-induced supercirculation effects from thrust vectoring have indicated a potential for not only increasing maneuverability of fighter aircraft but also as a means of improving cruise performance. The current study investigated a partial-span rectangular jet-exhaust nozzle located at the wing trailing edge that acts similar to a jet flap by increasing lift due to supercirculation. This paper summarizes experimental studies including the effects of nozzle deflection angle, wing camber, and nozzle shape and exit location on lift, drag and load distributions. The results indicate that significant increases in thrust-induced lift along with substantial decreases in drag are possible.

Evaluation of Reissner's Correction for Finite Span Aerodynamic Effects

The first term on the right-hand side of Eq. (14) is recognized as being the vortex distribution of a parabolic arc meanline which is symmetric about the midchord, and whose ideal angle of attack therefore is zero. The second term on the right-hand side is the vortex distribution of a flat plate airfoil at angle of attack. The constant C determines the lift lag of the additional distribution in unsteady flow. This constant can be found by applying the boundary condition at the airfoil surface, making use of the Cauchy-Riemann condition, and performing an integration. This will not be done here, since it is believed that the lag is insignificant for all cases of practical interest to maximum lift. Note that there is no lift lag for the effective camber change of the airfoil. The effective camber rjc may be found by substituting Eq. (14) into the following expression from thin airfoil theory, and integrating.

Transonic buffet characteristics of a 60-deg swept wing with design variations

Wind-tunnel tests of a typical swept wing fighter aircraft were conducted to determine the improvement in the buffet-onset lift coefficient CL^, resulting from wing camber and from leading and trailing edge deflection. The coefficient CLB was determined from trailing edge pressure, wing-root bending moment, fluorescent oil flow photographs, and force data. All four types of data are arranged in a composite graphical form to insure proper CLB selection and to identify the location on the wing where separation first occurs. Results at Mach 0.9 showed that 0.2 increase in camber (airfoil design lift coefficient) increased CLB approximately 0.08; a 5° leading edge droop 0.15; and a 5° inboard trailing edge flap 0.12. Analysis of design trends shown by previous experimental results are described for the effects of wing thickness ratio, sweep, aspect ratio, and camber on CLBSpecific buffet problems related to variable swept wings are discussed.

Effects of speed and camber on the ride behavior of a rapid transit vehicle traversing a dynamic roadway

Effects of speed and camber on the ride behavior of a rapid transit vehicle traversing a dynamic roadway