Chordwise Position Research Articles

Performance and control optimisations using the adaptive torsion wing

AbstractThis paper presents the Adaptive Torsion Wing (ATW) concept and performs two multidisciplinary design optimisation (MDO) studies by employing this novel concept across the wing of a representative UAV. The ATW concept varies the torsional stiffness of a two-spar wingbox by changing the enclosed area through the relative chordwise positions of the front and rear spar webs. The first study investigates the use of the ATW concept to improve the aerodynamic efficiency (lift-to-drag ratio) of the UAV. In contrast, the second study investigates the use of the concept to replace conventional ailerons and provide roll control. In both studies, the semi-span of the wing is split into five equal partitions and the concept is employed in each of them. The partitions are connected through thick ribs that allow the spar webs of each partition to translate independently of the webs of adjacent partitions and maintain a continuous load path across the wing span. An MDO suite consisting of a Genetic Algorithm (GA) optimiser coupled with a high-end low-fidelity aero-structural model was developed and employed in this paper.

Investigation of micro vortex generators on controlling flow separation over SCCH high-lift configuration

For the problem that the flow separation on the flap lowers the aerodynamic performance of high lift system, an investigation was carried out on using micro vortex generators (VGs) to control the separation on flap of the swept constant chord half-model (SCCH) high-lift configuration, at a small to medium angle of attack, by experimental and numerical methods. The basic flow characteristics of SCCH landing configuration were analyzed by using numerical method to provide required information for the design of micro VGs. Then, by keeping the cruise configuration intact, the preliminary design procedure and design methods of micro VGs were established. In addition, the micro VGs were designed. The effects of VG’s arrangement and geometric parameters, such as the arrangement mode, chordwise position, arrangement angle, height and spanwise distance, on controlling efficiency were investigated by using numerical method. Then the parameters of preliminary VGs were adjusted as the basis configuration for wind tunnel test. The experiments were accomplished in NH-2 wind tunnel for validating the numerical method, as well as obtaining the design principles and methods of micro VGs. The parameters of VGs were also optimized based on the experiments. The experimental results showed that the numerical design method can serve as an efficient and accurate design tool. The lift and drag were increased by 10% and 14%, respectively in landing state, which satisfied the requirements for landing. Finally, it was concluded that the established design principles and methods for micro VGs in this investigation can be used in engineering application.

Aeroelastic modeling and analysis of the wing/engine system of a large aircraft

An aeroelastic model of the wing/engine system of a large commercial aircraft is established. Taking into account the engine inertia force and thrust, static aeroelastic deformation of the wing structure and load distributions including shear force, bending moment and torque are studied. Models of the clean wing and the wing/engine system are compared and effects of chordwise and spanwise locations of the wing-mounted engines on the wing's natural vibration and flutter characteristics are analyzed. The research indicates that the bending moment at the root part of the wing is alleviated by the engine, and the combination of the engine inertia force and thrust will change the torque there remarkably, while not affect the bending and torsion deformations of the wing structure much. The results also show that wing-mounted engines will affect the wing's natural frequency of vibration, consequently having great effects on the flutter characteristics of the wing. Besides the classic wing flutter induced by the coupling of bending and torsion modes, a low-damping flutter type appears from the coupling of engine-pitch mode and the bending and torsion modes of the wing. Variations in chordwise and spanwise position of the wing-mounted engines will change the frequency of the engine-pitch mode, and further affect the low-damping flutter characteristics mentioned above, to which much attention should be paid in engineering application.

Bubble Burst Control Using Smart Structure Sensor Actuators for Stall Suppression

Based on our previous experimental work conducted pertinent to the flow separation bubble burst control using a thin plate (hereafter burst control plate) attached to the leading-edge of laminar airfoils, the burst control plate in the current study is incorporated in the smart structure sensor actuators system on a NACA 631-012 airfoil for stall suppression, lift enhancement and drag reduction. The smart structure sensor actuators system uses a pressure transducer, a displacement sensor and a feedback controller to operate and to monitor the actuators in an autonomous mode. Note that the height of the burst control plate depends on the angular movement of the actuators located on either end of the airfoil section. Experimental studies were conducted at two different Reynolds numbers of 1.3 x 10 5 and 1.9 x 10 5 based on the airfoil chord length. The wind tunnel experiments demonstrated the effectiveness of the smart structure sensor actuators system and the burst control plate located at 6% chord-wise position (defined at the trailing-edge of the plate) of the airfoil, in that the drag is reduced, stall angle is delayed, and a significant improvement in the lift-to-drag ratio is achieved over a wide range of angles-of-attack. The feedback control approach provides a means to strengthen the vortical structures inside the separated shear layer with the overall drag of the airfoil is kept minimal. The experimental results demonstrated the application of the smart structure sensor actuators system and the burst control plate on the NACA 631-012 airfoil can be effective means of bubble burst control and airfoil stall suppression in low speed air vehicle applications.

Lagwise Loads Analysis of a Rotor Blade with an Embedded Chordwise Absorber

An aeroelastic simulation of a stiff in-plane hingeless rotor in forward flight is conducted to evaluate the feasibility of reducing lagwise root loads via an embedded absorber within the blade. A finite element analysis based on a moderate deflection beam model is employed to capture the flap, lag, and torsion deflections of the rotor blade. The mass-spring-damper absorber system is simulated as a one-degree-of-freedom rigid body with chordwise motion. By using the Hamilton principle, system equations of motion are derived based on the generalized force formulation. Dynamic responses are calculated to investigate the potential of the absorber for in-plane loads reduction. Parametric studies are conducted to explore the influence of the absorber on the dynamic characteristics of the rotor blade. From the aeroelastic analysis of the blade-absorber system, results indicate that placing an embedded absorber in the blade chordwise direction can decrease the 1 and 2/rev lagwise root bending moments by about 50 and 90%, respectively, under large load states. The effects of absorber frequency, absorber damping, chordwise position of the absorber, rotor thrust, forward speed, and trim condition are also studied.

Application of Piezoelectric Actuators at Subsonic Speeds

An experimental study has been conducted to investigate the effects of using a series of piezoelectric actuators at two chordwise positions (25 and 50% chord length) on a NACA 0015 aerofoil at subsonic speeds. The lift and drag forces were measured using a three-component force balance. A detailed boundary-layer survey was carried out using hot-wire anemometry. The actuators operating at 25 % chord length are more effective, resulting in an increase of the lift-to-drag ratios at higher frequencies. Boundary-layer surveys indicated that, with the actuators operating at 25% chord length, the boundary-layer thickness is drastically reduced downstream of their location, whereas the flow remains mostly undisturbed in the upstream direction. The actuators can fix laminar to turbulent transition and eliminate the laminar separation bubble at lower Reynolds numbers.

Experimental Studies on Unsteady Lateral Blowing on NACA 0012

T HE initial study of steady lateral blowing was conducted by Dixon in 1969 [1]. The mechanism of this concept is analogous to the flow over slender delta wing platforms, which produces a leading-edge vortex and nonlinear lift curves. Steady lateral blowing might be thought of as providing the sweep effects such as those of the delta wings, for which the effective sweep is a function of the jet momentum. The lateral blowing jet can be described in terms of the chordwise position along the airfoil section, the nozzle height, and the nozzle size. Dixon et al. [2] andClarke [3] suggested that the best vertical nozzle positionmay be a function of the nozzle diameter and that the jet from a nozzle placed too near the airfoil surface may have a deleterious effect on the flow over the airfoil. Wong and Kontis [4] performed a comprehensive study on the steady spanwise blowing on a NACA 0012 airfoil section. The force measurements showed that the lift coefficient increasedwhen the steady lateral blowingwas applied, and the most effective blowing location was found to be at x=c 0:25 (x is the chordwise location and c is the airfoil chord), because the lift augmentation ratio Cl=C was always above one for all positive angles of attack [ Cl is the increment in the lift coefficient due to blowing, andC is the blowing momentum coefficient; the definition ofC is shown inEq. (1)].However, largemomentumwas required to generate the lift enhancement on the airfoil and this became a major issue in the effectiveness of blowing. As a consequence, Meyer and Seginer [5] performed some initial tests for generating the same lift increase using lower momentum through pulsed blowing. One of the main disadvantages of lateral blowing is due to the fact that air for blowing is bled from the compressor stage of the engine. This reduces the amount of air for combustion, hence reducing thrust. Unsteady blowing provides the beneficial effect of reducing the mass-flow requirement from the engine. Another adverse effect of the technology arises from the need to install plumbing, control valves, regulators, oscillators, etc., to supply and control the compressed air from the engine. This adds weight to the aircraft, and a higher lift is required. C thrust

Aeroelastic instability problems for wind turbines

AbstractThis paper deals with the aeroelastic instabilities that have occurred and may still occur for modern commercial wind turbines: stall‐induced vibrations for stall‐turbines, and classical flutter for pitch‐regulated turbines. A review of previous works is combined with derivations of analytical stability limits for typical blade sections that show the fundamental mechanisms of these instabilities. The risk of stall‐induced vibrations is mainly related to blade airfoil characteristics, effective direction of blade vibrations and structural damping; whereas the blade tip speed, torsional blade stiffness and chordwise position of the center of gravity along the blades are the main parameters for flutter. These instability characteristics are exemplified by aeroelastic stability analyses of different wind turbines. The review of each aeroelastic instability ends with a list of current research issues that represent unsolved aeroelastic instability problems for wind turbines. Copyright © 2007 John Wiley & Sons, Ltd.

Aerodynamics of Tip Leakage Flows Near Partial Squealer Rims in an Axial Flow Turbine Stage

This paper deals with an experimental investigation of aerodynamic characteristics of full and partial-length squealer rims in a turbine stage. Full and partial-length squealer rims are investigated separately on the pressure side and on the suction side in the “Axial Flow Turbine Research Facility” (AFTRF) of the Pennsylvania State University. The streamwise length of these “partial squealer tips” and their chordwise position are varied to find an optimal aerodynamic tip configuration. The optimal configuration in this cold turbine study is defined as the one that is minimizing the stage exit total pressure defect in the tip vortex dominated zone. A new “channel arrangement” diverting some of the leakage flow into the trailing edge zone is also studied. Current results indicate that the use of “partial squealer rims” in axial flow turbines can positively affect the local aerodynamic field by weakening the tip leakage vortex. Results also show that the suction side partial squealers are aerodynamically superior to the pressure side squealers and the channel arrangement. The suction side partial squealers are capable of reducing the stage exit total pressure defect associated with the tip leakage flow to a significant degree.

Control of Low-Pressure Turbine Separation Using Vortex-Generator Jets

The application of vortex-generator jets to control separation on the suction surface of a low-pressure turbine blade is reported. Blade Reynolds numbers in the experimental, linear turbine cascade match those for high-altitude operation of many aircraft gas-turbine engines, as well as the last stages of industrial ground-based gas turbines. Results are presented for steady blowing at jet blowing ratios from zero to four and at several chordwise positions and two freestream turbulence levels. Findings show that above a minimum blowing ratio, which is dependant on the injection location, the pressure loss in the modified blade's wake is reduced by a factor of between two and three. Boundary-layer traverses show that separation is almost completely eliminated with the application of blowing. No significant deleterious effects of vortex-generator jets are observed at higher (nonseparating) Reynolds numbers. The addition of 4% freestream turbulence to the cascade freestream lowers the separation Reynolds number of the turbine blade studied, but does not eliminate the effectiveness of the control technique. The vortex-generator jet control strategy is demonstrated to be a viable technique for low-pressure turbine separation control.

ボルテックス・フラップ付きデルタ翼の静的ロール特性について

Static normal force and rolling moment measurements for 60° and 70° delta wings with vortex flaps have been done for different roll angles at a Reynolds number based on a centerline chord of 1.3×105. Smoke visualization tests have also been done. Results show that the 60° delta wing is statically unstable in roll at the angle of attack of 35° when vortex flaps were deflected 30° downward. This is because the chordwise position where the vortex breakdown occurs is different for the left and right wings when in roll. The 70° delta wing with vortex flaps possesses static roll stability. Static rolling moment hysteresis has been observed for the 60° delta wing with a flap deflection angle of 30° at an angle of attack of 45°.

Turbine Separation Control Using Pulsed Vortex Generator Jets

The application of pulsed vortex generator jets to control separation on the suction surface of a low-pressure turbine blade is reported. Blade Reynolds numbers in the experimental, linear turbine cascade match those for high-altitude aircraft engines and aft stages of industrial turbine engines with elevated turbine inlet temperatures. The vortex generator jets have a 30 deg pitch and a 90 deg skew to the free-stream direction. Jet flow oscillations up to 100 Hz are produced using a high-frequency solenoid feed valve. Results are compared to steady blowing at jet blowing ratios less than 4 and at two chordwise positions upstream of the nominal separation zone. Results show that pulsed vortex generator jets produce a bulk flow effect comparable to that of steady jets with an order of magnitude less massflow. Boundary layer traverses and blade static pressure distributions show that separation is almost completely eliminated with the application of unsteady blowing. Reductions of over 50 percent in the wake loss profile of the controlled blade were measured. Experimental evidence suggests that the mechanism for unsteady control lies in the starting and ending transitions of the pulsing cycle rather than the injected jet stream itself. Boundary layer spectra support this conclusion and highlight significant differences between the steady and unsteady control techniques. The pulsed vortex generator jets are effective at both chordwise injection locations tested (45 and 63 percent axial chord) covering a substantial portion of the blade suction surface. This insensitivity to injection location bodes well for practical application of pulsed VGJ control where the separation location may not be accurately known a priori.

1998 Turbomachinery Committee Best Paper Award: An Experimental Study of Tip Clearance Flow in a Radial Inflow Turbine

This paper describes an experimental investigation of tip clearance flow in a radial inflow turbine. Flow visualization and static pressure measurements were performed. These were combined with hot-wire traverses into the tip gap. The experimental data indicate that the tip clearance flow in a radial turbine can be divided into three regions. The first region is located at the rotor inlet, where the influence of relative casing motion dominates the flow over the tip. The second region is located toward midchord, where the effect of relative casing motion is weakened. Finally, a third region exists in the exducer, where the effect of relative casing motion becomes small and the leakage flow resembles the tip flow behavior in an axial turbine. Integration of the velocity profiles showed that there is little tip leakage in the first part of the rotor because of the effect of scraping. It was found that the bulk of tip leakage flow in a radial turbine passes through the exducer. The mass flow rate, measured at four chordwise positions, was compared with a standard axial turbine tip leakage model. The result revealed the need for a model suited to radial turbines. The hot-wire measurements also indicated a higher tip gap loss in the exducer of the radial turbine. This explains why the stage efficiency of a radial inflow turbine is more affected by increasing the radial clearance than by increasing the axial clearance.

On the Absolute Instability of the Attachment-Line and Swept-Hiemenz Boundary Layers

RID=""ID="" Communicated by M.Y. HussainiAbstract: Recently it has been shown that flow over a rotating-disk is absolutely unstable. In this paper we investigate the absolute instability of the related swept-Hiemenz and attachment-line boundary-layer flows. The linearized stability equations are obtained and the eigenvalues of the dispersion relation are found by solving the full stability equations in Fourier-transform space using a spectral method. Unlike previous work on this problem, no quasi-parallel approximation has been made and all the terms appearing in the stability equations have been retained. We were unable to locate branch points satisfying the Briggs–Bers criterion for the attachment-line boundary layer suggesting that this flow is only convectively unstable. However, for the swept-Hiemenz boundary layer our results show that this flow becomes absolutely unstable (in the chordwise direction), starting from the leading-edge extending up to a chordwise position of approximately 310 for some particular spanwise Reynolds numbers. It is found that the retention of all the terms in the full system of equations leads to results which are more unstable, in terms of absolute instability, than the Orr–Sommerfeld system studied by others. The techniques used here apply equally to other non-parallel two- and three-dimensional boundary-layer flows.

White muscle strain in the common carp and red to white muscle gearing ratios in fish

White muscle strains were recorded using sonomicrometry techniques for 70 fast-starts in the common carp Cyprinus carpio L. High-speed cine images were recorded simultaneously for 54 of these starts, and muscle strain was calculated independently from the digitized outlines of the fish. Sonomicrometry measurements of superficial muscle strain were not significantly different from the strain as calculated from the theory of simple bending of a homogeneous material: superficial muscle strain thus varied with chordwise distance from the spine. However, white muscle strain across a transverse section of the myotome shows less variation with chordwise position than would be expected from simple bending theory. Muscle strains measured using sonomicrometry thus do not necessarily represent the more uniform strain predicted for the whole section of the fish. White muscle strain can be accurately predicted from the spine curvatures as measured from the cine images if the gearing ratio between the red and white muscle fibres is known. A model for calculating the gearing ratio from the helical muscle fibre geometry was re-evaluated using current data for the kinematics of fast-starting C. carpio. This model predicted a mean gearing ratio of 2.8 for these fast-starts. A quicker, alternative approach to estimating gearing ratio from the position of the centroid of white fibre area is proposed and results in ratios similar to those calculated from the model of helical geometry. White muscle strains in fish can thus be estimated from measurements of spine curvature and muscle distribution alone.

Studies of Secondary Flow at Endwall of a Supersonic Compressor Cascade

A combined experimental and numerical study was conducted to investigate the endwall secondary flow in a linear supersonic compressor cascade. The experiment was carried out in a supersonic wind tunnel, with Mach and Reynolds numbers (based on the blade chord) of 2.36 and 4.8 X 106, respectively. A three-dimensional viscous computational fluid dynamics (CFD) calculation was performed to simulate the endwall boundary layer and the three-dimensional secondary flows. The experimental and computational analyses presented are all at the design incidence angle. Experimental measurements include multiple-color surface oil flow visualization applied to the endwall and blade surfaces, blade-to-blade pitot pressure contours at the 90% chordwise position, and upstream and downstream static pressures at the endwall. All of these measurements compare favorably with the CFD calculation. Results from the CFD analysis show that endwall total-pressure losses in this supersonic compressor cascade are a significant contributor to the overall losses and can account for one-third of the overall flowfield loss.

Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part II—Impingement Cooling With Film Coolant Extraction

An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions that allow extraction, or bleed-off, of a portion of the internal cooling flow to provide leading edge film cooling along the blade external surface. The internal impingement air is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and generally exiting, minus the bleed-off flow, in the opposite or chordwise direction. The film coolant flow extraction takes place through two lines of holes, one each on the blade suction side and the blade pressure side, both fairly close to the airfoil leading edge. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surfaces, and are observed during a test transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice is covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise-average Nusselt numbers and spanwise Nusselt number gradients as functions of chordwise position. The results without film coolant extraction, presented in the Part I companion paper, are used as a basis of comparison to determine the additional effects of the film cooling bleed. Results indicate that heat transfer is primarily dependent on jet Reynolds number with smaller influences from the flow extraction rate. The results also suggest that changes in the spanwise alignment of the impingement nozzles relative to the position of the film cooling holes can cause significant variations in leading edge metal temperatures.

Local Heat Transfer in Internally Cooled Turbine Airfoil Leading Edge Regions: Part I—Impingement Cooling Without Film Coolant Extraction

An experimental study has been designed and performed to measure very localized internal heat transfer characteristics in large-scale models of turbine blade impingement-cooled leading edge regions. Cooling is provided by a single line of equally spaced multiple jets, aimed at the leading edge apex and exiting the leading edge region in the opposite or chordwise direction. Detailed two-dimensional local surface Nusselt number distributions have been obtained through the use of aerodynamically steady but thermally transient tests employing temperature-indicating coatings. The thin coatings are sprayed directly on the test surface and are observed during the transient with automated computer vision and data acquisition systems. A wide range of parameter combinations of interest in cooled airfoil practice are covered in the test matrix, including combinations of variations in jet Reynolds number, airfoil leading edge sharpness, jet pitch-to-diameter ratio, and jet nozzle-to-apex travel distance. Measured local Nusselt numbers at each chordwise location back from the stagnation line have been used to calculate both the spanwise average Nusselt number and spanwise Nusselt number gradient as functions of chordwise position. Results indicate general increases in heat transfer with approximately the 0.6 power of jet Reynolds number, increases in heat transfer with decreasing leading edge sharpness as well as with decreasing nozzle-to-apex distance, and increases in spanwise average heat transfer with decreasing jet pitch-to-diameter ratio. The latter increases are accompanied by increases in the spanwise gradient of the heat transfer coefficient. Comparison with available prior results of much coarser spatial resolution shows good agreement and establishes confidence in the use of the results for design purposes and as baseline results for comparison with subsequent experiments involving film cooling bleed.

Boundary-Layer Development on an Axial-Flow Compressor Stator Blade

The boundary layer on an axial-flow compressor stator blade has been measured using an ensemble-averaging technique. Although the mean velocity profiles appear to indicate fully developed turbulent flow, ensemble-averaged instantaneous profiles show the boundary layer to be highly unsteady and transitional over much of the blade chord. At a given chordwise position, variations in boundary-layer thickness with time of up to 150 percent were recorded. When compared to boundary-layer development on a similar blade in a two-dimensional cascade the stator blade boundary-layer growth was found to be much greater. The results indicate that extreme caution should be used in attempting to predict blade boundary-layer development from cascade test results or steady calculation procedures.

Unsteady Kutta condition at high values of the reduced frequency parameter

The unsteady Kutta condition is discussed in the light of some recent experimental measurements made near the trailing edge of a long flat plate and a 10C4 airfoil. The hierachy of disagreement from the theoretically predicted zero trailing edge loading caused by viscous instabilities is found to be acoustically correlated vortex shedding, natural vortex shedding, Tollmien-Schlichting waves, and, by implication, turbulent boundary-layer eddies. The region of significant chordwise disagreement scales with the wake perturbation wavelength of the corresponding instability. Coordinating the vorticity of the turbulent boundary layer shed from the profile airfoil with a transverse acoustic resonance produced a distinct disagreement of the Kutta condition at high reduced frequency parameters ( = wc/U). In this case and for vortex shedding, the extrapolated loading coefficient at the trailing edge increased with the nondimensional acoustic amplitude. I. Introduction T^HE Kutta condition as applied in unsteady JL potential airfoil analyses is essentially an extension of steady theory. Kutta ! postulated that a value of circulation should be chosen in his steady potential model to avoid a velocity singularity at the sharp trailing edge of an airfoil. This condition can be established if the trailing edge is also the rear stagnation point. The resulting modeled flow pattern agrees with that observed in steady flow and also predicts the lift and its chordwise distribution well at low angles of attack. The theoretical consequences of this hypothesis are that the lift loading or chordwise vorticity jump approaches zero at the trailing edge. An alternative statement is that the surface velocities on either side of the airfoil approach a common value at the rear stagnation point. For rounded trailing edges, the position of the rear stagnation point is indeterminate , as there is no velocity singularity to be avoided and so fix its location. In this case and for the situation of real flows with viscosity, Taylor2 proposed the condition of zero net vorticity discharge to establish the steady lift value. Preston3 explained the deviation of the lift of an airfoil at low angles of incidence from the potential theory value as due to the profile alteration from the boundary-layer growth. His calculations incorporated Taylor's vorticity discharge condition. Various approximate steady lift calculation methods for the rounded trailing edge geometry have been proposed by Gostello 4 and others. These extend the upper and lower lift distributions, at a selected chordwise position, to the trailing edge to give zero loading and thereby remove the stagnation point indeterminacy. In the unsteady case there are all the previous theoretical difficulties and, in addition, the unsteady effects on the viscous boundary layer and the shed vorticity. The latter complicates the airfoil response, making it a function of the airfoil's vorticity history. However, same theoretical assumption for the Kutta condition, of no unsteady loading at the