Pitching moment

About: Pitching moment is a research topic. Over the lifetime, 3213 publications have been published within this topic receiving 38721 citations.

Experimental Results Analysis for UiTM BWB B aseline!I and Baseline!II UAV Running at 0.1 Mach number

Abstract: This paper presents results analysis for two models of UiTM Blended Wing Body (BWB) UAV tested in UiTM Low Speed Wind Tunnel. The first model is known as the BWB Baseline!I and the new model known as BWB Baseline!II. The Baseline!II has a simpler planform, broader!chord wing and slimmer body compared to its predecessor while maintaining wingspan. The wind!tunnel experiments were executed at around 0.1 Mach number or about 35m/s with 1/6 scaled down model. Baseline!I is designed with centre elevator while Baseline!II uses canard for pitching motion purpose. The experiments were carried out at various elevator and canard deflection angles. The lift coefficient, drag coefficient, pitching moment coefficient, L/D ratio and drag polar curves were plotted to show the performance of aircraft at various angle of attack. For zero elevator and canard deflection the results show similar trends in terms of lift curve, drag curve and pitching moment curves for both aircrafts.

Improved Potential Gradient Method to Calculate Airloads on Oscillating Supersonic Interfering Surfaces

Abstract: sectional normal force coefficient in AGARD notation, Eq. (7) = sectional pitching moment about quarter-chord point coefficient in AGARD notation, Eq. (8), C*mi is moment about midchord = variation of the potential jump behind the leading edge of an element on which a constant potential gradient in streamwise direction is assumed (Fig. Ib) = amplitude of the normal displacement of the oscillating wing = reduced frequency = ul/ U supersonic pressure dipole kernel = reference length = Mach number = normal directions of receiving wings = total number of unknowns = dynamic pressure = /2pU generalized aerodynamic coefficients , generalized forces, Eq. (6)

Measured forces and moments on a delta wing during pitch-up

Abstract: A series of low-speed wind tunnel tests on a 70-deg, sharp, leading-edge delta wing undergoing ramp pitching motion of high amplitude were performed to investigate the aerodynamic forces and moments. Forces and moments were obtained from a six-component interanl balance. Large amplitude oscillatory motion was produced by sinusoidally oscillating the model over a range of reduced frequencies. Ramp motion was produced by pitching the model through a half cycle of sinusoidal motion at a root chord Reynolds number of 1.54 million. The effect of ramp and oscillatory motions on the forces and moments are almost identical at matched pitch rates. Pitch rate had strong effect on the magnitude of the aerodynamic forces and moments. Upon completion of the model motion, some time is required for the forces and moments to decay to their static values. This convergence of the dynamic values to the static ones was a function of the pitch rate.

Unsteady Airfoil with Dynamic Leading- and Trailing-Edge Flaps

Abstract: T HE dynamic overshoot in lift force and the accompanying large nose-down pitching moment observed on oscillating and pitching airfoils, as a result of the formation, convection, and shedding of an energetic dynamic-stall vortex (DSV), continue to make dynamic stall and its control an important topic in unsteady aerodynamics. Moreover, once the DSV passes the airfoil trailing edge and moves into the wake, the flow progresses to a state of poststall full separation over the upper surface and an abrupt loss of lift is incurred, causing a large degree of hysteresis in the lift coefficient. Various dynamic-stall flow control methods, such as trailing-edge flaps [1–5], pulsating and synthetic jets [6–8], leadingedge slots with and without blowing and suction [9,10], and dynamically deformable and variably drooped leading edges [11,12], etc., capable of minimizing, or eliminating, the hysteresis in the dynamic-Cl loop and the peak negative pitching-moment coefficientCm;peak, have been proposed. It should, however, be noted that for rotorcraft, the dynamic-stall flow control is aimed at the prevention of the DSV occurrence and the mitigation of nose-down pitching moment on rotor blades, whereas for highly maneuverable aircraft, the purpose is to delay the DSV spillage and to enhance the dynamic lift. The trailing-edge flap (TEF) dynamic-flow control concept has been considered widely for dynamic lift enhancement and nose-down pitching-moment suppression. Trailing-edge flaps have also been used extensively as a routine practice of controlling the lift by temporarily altering airfoil camber on an airplane in steadylow-speed operations, especially during takeoff and landing, without penalizing cruise performance. Rennie and Jumper [2] investigated the effectiveness of deflecting a 27% TEF control surface in negating the unsteady lift generated during a pitchmotion of a NACA 0009 airfoil atRe 2 10. They reported that the dynamic TEF effectiveness was larger than the steady-state value and had direct ramifications on unsteady lift control. The effectiveness of a deflecting control surface was also found to be higher while the control surface was in motion and was rapidly decreased at large deflections, as a result of the thickening or separation of the boundary layer on the trailing-edge flap. Furthermore, the motion of TEF appeared to delay the occurrence of dynamic stall on the airfoil. Gerontakos and Lee [5] examined the alleviation of the nose-downpitchingmoment by using a 25%c TEF, actuated dynamically in response to the airfoil oscillation, at Re 1:65 10. The prescheduled TEF motion consisted of a brief pulse, represented by a constant ramp-upmotion, remained steady briefly, and was followed by a constant ramp-down motion. They found that the reduction in j Cm;peakj was mainly a consequence of the suction pressure introduced on the lower surface of the upward deflected TEF and that a relatively early flap actuation (initiated between the static-stall angle ss and themaximum angle of attack max during pitch-up) with a rather long duration (about half of the oscillation cycle time) and a maximum deflection max 60% max should be more effective at reducing the nose-down pitching-moment excursion, and, in the meantime, provide a good compromise between the various aerodynamic requirements. Also, the magnitude of the maximum lift coefficientCl;max was found to be rather insensitive to the flap-actuation duration td, whereas the poststall lift was decreased with decreasing td. More important, the DSV formation and detachment were not affected by the upward TEF motion. In addition, dynamic leading-edge-flap lift control has also been employed by researchers elsewhere to allow high-performance aircraft to achieve a rapid and sustained high-lift coefficient. The purpose of leading-edge devices is to increase camber and thus suppress leading-edge separation during rapid arbitrary airfoil pitching maneuvers. Rennie and Jumper [13] reported an experimental determination of a 20% leading-edge flap schedule used to maintain attached flow during arbitrary dynamic pitching motions of a NACA 0009 airfoil at Re 2 10. The leading-edge flap (LEF) schedule kept the flow attached dynamically at angles of attack that were separated during static tests. The use of the LEF avoided catastrophic dynamic-stall flow-separation events. Also, for a static leading-edge flap schedule, the attached flow could be maintained by keeping the leading-edge flap aligned with the oncoming flow. In summary, it is known that both leadingand trailing-edge flaps can serve to increase the camber and manage the boundary layer effectively and thus increase the maximum lift of an airfoil. However, despite much predictive work, published experimental data on the unsteady aerodynamic loads induced by the dynamically deflecting leadingand trailing-edge flaps are still sparse. A preliminary experimental study was conducted to examine the effects of an 18%c leadingand a 25%c trailing-edge flap, actuated dynamically and independently, on the dynamic-load loops of a sinusoidally oscillatingNACA0015 airfoil in a subsonicwind tunnel at Re 2:86 10. The dynamic-load loops were obtained by integrating the unsteady surface pressure distributions. Both upward and downward flap deflections actuated at a fixed start time (ts 0 ) were investigated. This ts value was chosen to maximize the influence of the flap motion on the transient DSV-induced effects. Special emphasis was also placed on the simultaneousmeasurements of airfoil and flap deflection histories, synchronized with the surface pressure measurements.

Aerodynamic Analysis of Grid-Fin Configurations at Supersonic Speeds

Abstract: An aerodynamic prediction method for the estimation of aerodynamic characteristics of grid-fin configurations at supersonic Mach numbers has been developed. The method is based on shock-expansion theory. The shock and expansion relations and the interactions between the shock and expansion waves are used to predict the pressure distribution inside each grid-fin cell. The effect of body is modeled using doublet in the crossflow plane. The effect of symmetric separated vortices from the leeward side of body is also modeled using empirical data for the strength and location of vortices. The present method has been validated with available experimental data for body grid-fin configurations at angles of attack without and with fin deflection. The comparison of normal force coefficient on the individual grid fins as well as the overall normal force, pitching moment, and axial force coefficient with experimental data is good. Effect of roll orientation on normal force, pitching moment, and induced out-of-plane force is also brought out.