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The efficiency of boundary-layer ingestion (BLI) propulsion systems is closely related to the amount of ingested boundary layer. A high BLI fraction requires low propulsor flow power. In this paper, the internal flow characteristics and performance of an S-shaped inlet with a high BLI, up to 83.8% of the inlet capture height, are explored in detail under a freestream Mach number of 0.75. The results reveal that the middle and lower parts of the aerodynamic interface plane (AIP) section are packed with low-energy flows dominated by a pair of vortices, formed by the coalescence of side and separation vortices. At MAIP=0.46, the side vortex, originating near the first inlet bend, is caused by a spanwise pressure gradient, while the side vortex is connected to the horseshoe vortex at the root of the inlet lip at MAIP=0.33. The vortex coalescence results in a sudden increase in the strength of the vortex core and a stepwise increase in circulation (ΓRx). The flow vortex intensity in the inlet is divided into three stages from the distribution of ΓRx. In addition, a reduction in MAIP increases the intensity of secondary flow along the section, particularly the intensity at the flow separation zone. With increasing MAIP, the swirl intensity decreases, whereas the total pressure distortion index DC60 rises, reaching 0.54, which is beyond the requirement of turbomachinery component flow uniformity. [ABSTRACT FROM AUTHOR]

A novel global parametric system identification framework is introduced in this work for aeroelastic modeling under varying flight states. The presented framework is based on the functionally pooled (FP) time-series models that enable explicit analytical inclusion of any admissible flight states into the model parameters and thus the system dynamics. In this paper, the autoregressive (AR) type of the FP model, which is designated as the FP-AR, is employed to interpret the aerodynamics of a wing structure. The data were recorded by accelerometers during a dedicated wind-tunnel flutter test with the airspeed increasing all the way to the flutter boundary. Including the varying flight states defined by the increasing airspeed into the system modeling via the FP technique, the global framework provides a more sophisticated identification results compared with the traditional nonparametric Welch-based spectral estimation. Flutter occurrence is indicated by the stability margin of the aeroelastic system evaluated by the estimated FP-AR parameters based on Jury's stability criterion. For online application purpose, an iterative state-based FP-AR process is also proposed in this paper. Compared with the standard FP-AR modeling, the iterative realization is a developing process that provides a global approximation of the system dynamics at each flight state. Experimental evaluation demonstrates the feasibility and effectiveness of the proposed global framework in aeroelastic modeling while facilitating an accurate flutter boundary prediction. [ABSTRACT FROM AUTHOR]

Compound helicopters often have to reduce their rotor tip speeds to achieve higher forward speeds. The rotors, therefore, have to undergo a transient process of rotor speed changes. To investigate the transient loads and the related transformation from the rotating frame to the nonrotating frame, a rotor model is used to predict the transient aeroelastic responses. The rotor model is validated by using test data of a teetering model rotor during engagement and disengagement operations. The investigations indicate that a small transient lagwise root bending moment can result in a large transient contribution to rotor torque. This transient component is triggered by the force related with the sudden change of the angular acceleration of the rotor speed. Increasing the blade damping in the lagwise direction can significantly decrease the transient components of the lagwise root bending moment and the corresponding rotor torque. Making the angular acceleration continuous will not trigger the transient components, suggesting that this is a better way for the reduction of the transient rotor loads during speed changes. [ABSTRACT FROM AUTHOR]

A key design factor impacting the use of electrical power to drive aircraft systems and subsystems is energy efficiency. With the design of an all-electric, hybrid ice protection system, energy consumption can be reduced to a large extent. The hybridization is achieved through an intentional partitioning of the ice at the stagnation line by melting via surface heating and ice shedding in the unheated regions of the airfoil surface via an electromechanical deicing system based on piezoelectric multilayer actuators. To further reduce energy consumption, the adhesion forces between the ice and the airfoil surface can be reduced using an ultrasmooth, nanostructured surface with water- and ice-repellent properties that encourages ice shedding. Experimental investigations, performed in a laboratory-scale icing wind tunnel for a smallscale configuration, reveal that the hybrid approach for ice protection reliably sheds the ice accreted on the airfoil surface. Compared with an all-thermoelectric system for ice protection investigated in the same icing wind tunnel facility using identical test conditions, the hybrid approach was demonstrated to reduce power consumption up to 91 %. Beyond the laboratory tests, numerical simulations of the hybrid strategy analogous to the one used for the experiments are performed. The time history of the residual ice shapes aft of the heated region is simulated using the ice accretion prediction software LEWICE2D for a wet-running anti-icing subsystem. Finite element analyses of the effects of the piezoelectric actuators are then performed using Abaqus to investigate the ice-shedding capability in the unheated regions of the airfoil surface. The numerical results show that the variation in the different ice shapes affects the stiffness of the model. It becomes obvious that the critical threshold for ice shedding, that is, the stiffness that determines whether residual ice delaminates from the airfoil surface, is affected to a minor extent by the interfacial area and predominantly by the thickness of the ice layer. Further, the simulation results correlate well with experimental results obtained in the icing wind tunnel. It can be concluded that reliable operation of the hybrid system for ice shedding can be guaranteed when using a harmonic sweep excitation able to excite the structure at its resonance. [ABSTRACT FROM AUTHOR]

The aim of this paper is to present aerodynamic analyses of realistic fuselage configurations using computational fluid dynamics. A development model of the Ansat helicopter is employed. The model is tested at the subsonic wind tunnel of the Kazan National Research Technical University /Kazan Aviation Institute for a range of Reynolds numbers, pitch, and yaw angles. The computational fluid dynamics results were found to be in fair agreement with the test data and revealed flow separation at the rear of the fuselage. Once confidence in the computational fluid dynamics method was established, further modifications were introduced to the fuselage model to demonstrate drag reduction via small shape changes. The contribution to the overall drag from each fuselage part and the interference drag between the main fuselage components were also investigated. [ABSTRACT FROM AUTHOR]

The objective of this work was to highlight modeling considerations for estimating the required static-pressure-error corrections using computational fluid dynamics. This work was performed to complement the U.S. Air Force Test Pilot School activities that used a T-38C aircraft with a previously untested Pitot-static system. Before the flight test, it is highly desirable to have an estimate of how any modifications will affect sensor performance, and to possess the capability to explain any discrepancies from previous test programs. Estimates of this kind are only obtainable via wind-tunnel tests or through modeling and simulation. This work presents the computational fluid dynamics requirements and methodologies needed to estimate the static-pressure-error correction of a T-38C in the subsonic, transonic, and low supersonic regimes. Estimates are validated with flight-test results, and proposed solutions for further reduction of static-pressure error via modifications to the flight-tested Pitot-static-probe geometry are included. [ABSTRACT FROM AUTHOR]

Contributions of the unstructured Reynolds-averaged Navier-Stokes code FUN3D to the 2nd AIAA Computational Fluid Dynamics High-Lift Prediction Workshop are described, and detailed comparisons are made with experimental data. Using workshop-supplied grids, FUN3D results for the simplified high-lift configuration are compared with results from the structured code CFL3D. Using the same turbulence model, both codes compare reasonably well in terms of total forces and moment, and the maximum lift is similarly over-predicted for both codes compared to experiment. By including more representative geometry features such as slat and flap support brackets and slat pressure tube bundles, FUN3D captures the general effects of the Reynolds number variation, but under-predicts maximum lift on workshop-supplied grids in comparison with the experimental data due to excessive separation. However, when output-based, off-body grid adaptation in FUN3D is employed, results improve considerably. In particular, when the geometry includes both brackets and the pressure tube bundles, grid adaptation results in a more accurate prediction of lift near stall in comparison with the wind-tunnel data. Furthermore, a rotation-corrected turbulence model shows improved pressure predictions on the outboard span when using adapted grids. [ABSTRACT FROM AUTHOR]

The current approach used to apply uncertainty intervals to balance estimated loads is based on the root mean square error from calibration. Using the root mean square error, a constant interval is applied around the estimated load and it is expected that a predetermined percentage of the check-loads applied fall within this constant uncertainty interval. However, this approach ignores additional sources of uncertainty and assumes constant uncertainty regardless of the load combination and magnitude applied to the balance. Rigorous prediction interval theory permits varying interval widths but fails to account for the additional error sources that are unrelated to the mathematical modeling. An engineered solution is proposed that combines prediction interval theory and the need to account for the additional sources of uncertainty from calibration and check loading. Results from a case study using the in-situ load system show improved probabilistic behavior in terms of uncertainty interval capture percentage when compared with the current root mean square error method. [ABSTRACT FROM AUTHOR]

To investigate the effect of control-surface freeplay, an aeroelastic wind-tunnel model of a T-tail with freeplay in the control chain was developed. The T-tail rig presents a conventional vertical fin represented only by its main structural component: the main spar. No aerodynamic sectors have been used to reproduce the aerodynamic contribution of a horizontal tail. This was chosen because of the limited size of the wind-tunnel chamber. Two numerical models were developed. The first one describes the dynamics of the tail by a state-space system with nonlinearity represented as a lumped element in the actuator feedback loop, and the second is described by the high-order harmonic balance approach, where the response of the nonlinear system is approximated with a periodic signal. The two approaches are then validated against experimental measurements collected during wind-tunnel testing. The results of this validation process as well the numerical and experimental approaches adopted are reported in this paper. [ABSTRACT FROM AUTHOR]

A closed-loop control system has been developed for microtabs and demonstrated in the University of California, Davis, Aeronautical Wind Tunnel. Input to the system comes from two dynamic pressure transducers located at 15% chord on the upper and lower surfaces of the airfoil. Microtabs with a height of 1% chord are located at 95% chord on the lower surface and 90% chord on the upper surface, enabling rapid increase or decrease of the lift force through deployment of the tabs on the pressure or suction side of the airfoil, respectively. Tests have been conducted with variations of up to 23% in airspeed, and the control system is able to reduce the total deviation in the lift force by 60% compared to the performance of the airfoil without the control system. The experiment validates the effectiveness of closed-loop control of lift using a trailing-edge tab-based system for lifting surfaces such as wind-turbine rotor blades. [ABSTRACT FROM AUTHOR]

Results from the Facility Analysis Verification and Operational Reliability project revealed a critical gap in capability in ground-based aeronautics-research applications. Without a standardized process for check loading the wind-tunnel balance or the model system, the quality of the aerodynamic force data collected varied significantly between facilities. The In Situ Load System was developed to provide a standard for facilities in the check-loading process. The system includes both the hardware and a statistically rigorous process that facilitates the ability for the user to make defendable decisions on the performance of the system. The compactness and simplicity of the system reduce customer costs unrelated to achieving the research objectives, while simultaneously improving the knowledge about the accuracy of the test data collected. While the focus is on the check-load process, the hardware and methods are also applicable to the in situ calibration of a balance or wind-tunnel model system. [ABSTRACT FROM AUTHOR]

Maximum rotor lift capability is investigated using wind-tunnel test data of McHugh (modified 1/10-scale CH-47B rotor) and a full-scale UH-60A rotor. Rotor performance calculations with the comprehensive rotorcraft analysis CAMRAD n are compared with the wind-tunnel test data. The analysis of the McHugh rotor with the Reynolds-number-corrected airfoil table shows good correlation with the measurements for µ = 0.1 to 0.5 and is able to predict the maximum rotor lift reasonably well, especially at 0.2 ≤ µ 0.4. The analysis is also able to predict the maximum lift of the full-scale UH-60A rotor within about 3.5% at µ = 0.24 and 0.3. Calculations with dynamic stall models, in general, show only a small influence on the rotor performance and are not necessary to predict maximum lift. Airfoils have an important role in defining the maximum lift capability of the rotor. The VR-12 airfoil, which has stall characteristics superior to the baseline V23010 airfoil, substantially improves the maximum lift capability of the McHugh rotor, showing the potential to improve the behavior of a rotor by improving the airfoil's static stall characteristics. [ABSTRACT FROM AUTHOR]

Evaluation of the conventional linear wall-interference correction method is conducted using flows around a two-dimensional airfoil in a wind tunnel computed by computational fluid dynamics and a porous wall model. First, the computational results are validated by comparison with wind-tunnel test measurements. Then, the wall-interference effect is discussed using these results. Porous walls are effective for reducing blockage, which change the Mach number around the airfoil. However, high wall porosity causes large down wash, which changes the flow angle. In this paper, Mokry's method, which is one of the linear wall-interference correction methods, is evaluated quantitatively using computational fluid dynamics. Computational result with wind-tunnel walls are corrected by Mokry's method, and the corrected result is compared with the one without wind-tunnel walls. Mokry's method shows high accuracy in subsonic and no-stall conditions, which satisfy the requirement of the small-perturbation potential equation. Its accuracy is also high even if the model is large relative to the wind-tunnel size. However, an error of about 10 drag counts is caused due to a shock wave in transonic flow conditions, and accuracy rapidly degrades around the stall condition, where an error exceeding 100 counts is caused by flow separation. [ABSTRACT FROM AUTHOR]

Multivariate-orthogonal-function modeling was applied to wind-tunnel databases for eight different aircraft to identify a generic global aerodynamic model structure that could be used for any of the aircraft. For each aircraft database and each nondimensional aerodynamic coefficient, global models were identified from multivariate polynomials in the nondimensional states and controls, using an orthogonalization procedure. A predicted-square-error criterion was used to automatically select the model terms. Modeling terms selected in at least half of the analyses, which totaled 45 terms, were retained to form the generic global aerodynamic model structure. Least squares was used to estimate the model parameters and associated uncertainty that best fit the generic global aerodynamic model structure to each database. The result was a single generic aerodynamic model structure that could be used to accurately characterize the global aerodynamics for any of the eight aircraft, simply by changing the values of the model parameters. Nonlinear flight simulations were used to demonstrate that the generic global aerodynamic model accurately reproduces trim solutions, local dynamic behavior, and global dynamic behavior under large-amplitude excitation. This compact global aerodynamic model can decrease flight-computer memory requirements for implementing onboard fault detection or flight control systems, enable quick changes for conceptual aircraft models, and provide smooth analytical functional representations of the global aerodynamics for control and optimization applications. All information required to construct global aerodynamic models for nonlinear simulations of the eight aircraft is provided in this paper. [ABSTRACT FROM AUTHOR]

The S414, slotted, natural-Iaminar-flow airfoil was designed for rotorcraft applications and has been analyzed theoretically using MSES and OVERFLOW. In addition, it has been verified experimentally in the Pennsylvania State University Low-Speed, Low-Turbulence Wind Tunnel. The primary objectives of the design are high maximum lift and low profile drag, both of which have been achieved while satisfying a thickness constraint. The theoretical analyses show good agreement with the experimental results. The experimental results are compared with those of the S406 and S411 airfoils, which were designed to similar specifications, and illustrate the potential benefits of the slotted, natural-laminar-flow concept. [ABSTRACT FROM AUTHOR]

Targeted energy transfer is studied as a means for suppression of transonic aeroelastic instabilities of a wind-tunnel swept wing, with a focus on designing a lightweight nonlinear energy sink that improves the critical flutter condition. The aeroelastic response modes of the wing with a nonlinear energy sink coupled to the tip are identified and tested for robustness using a medium-fidelity computational aeroelasticity model, and confirm that robust suppression of transonic aeroelastic instabilities is achievable. Accordingly, a nonlinear energy sink is designed based on a parametric study, and its transonic aeroelastic effects are studied using medium- and high-fidelity models. The results of both models indicate an improvement in stability over a broad range of conditions; the high-fidelity model predicts an approximately 40% increase in the dynamic pressure at the critical stability condition. Finally, a prototype winglet-mounted nonlinear energy sink is modeled to examine its aeroelastic effects. The results show that the nonlinear-energy-sink design is robust, but the winglet design plays a critical role that must be considered in the overall effect. [ABSTRACT FROM AUTHOR]

Autorotation in a steady-state axial motion is investigated theoretically and experimentally. A simplified model presents the major physical parameters affecting this phenomenon, giving an insight into the phenomenon and the influence of various parameters. A more detailed model, based on a multibody dynamic approach, is also used to describe the phenomenon more accurately. A series of experiments of a rotor model in a wind tunnel is carried out using two kinds of rotor blades operating at a wide range of freestream velocities and pitch angles. The experimental results are used to verify the theoretical results obtained by the simplified and the detailed models. Good agreement between the calculated and measured results is shown. [ABSTRACT FROM AUTHOR]

A number of airfoils have been tested in the Pennsylvania State University low-speed low-turbulence wind tunnel, and the results of these tests are compared with those predicted using several well-known theoretical methods. The theoretical methods used are the potential-flow/integral boundary-layer methods XFOIL 6.96 and PROFIL07 and the Reynolds-averaged Navier-Stokes solver OVERFLOW 2.2e. This version of the OVERFLOW solver contains an implementation of the transitional shear-stress transport turbulence model developed by Langtry and Menter ("Correlation-Based Transition Modeling for Unstructured Parallelized Computational Fluid Dynamics Codes," AIAA Journal, Vol. 47, No. 12, 2009, pp. 2894-2906). This model is capable of capturing the influence of transition on the flowfield through a local-correlation method. The airfoils considered for this study are the E 387, $805, PSU 94-097, HTR1555, S903, and S824. Although none of the theoretical methods considered were consistently the best overall, all codes predicted the drag well in the low-drag regions of these airfoils and the codes incorporating integral boundary-layer methods generally agreed better with the experimental data. The methods also all showed inconsistencies in predicting the maximum lift coefficient, with XFOIL and OVERFLOW frequently overpredicting this value. [ABSTRACT FROM AUTHOR]

The article examines the accuracy of the typical light general-aviation angle-of-attack system is measured using a wind-tunnel testing tool located in the open-circuit induction at the U.S. Naval Academy. A probe is placed on the bottom of the wing of a single-engine aircraft or on the underside of the nose of a twin-engine aircraft in order to test the angle-of-attack, pitch angle and yaw angle for dynamic pressure.

Closed-loop active flow control is applied in wind-tunnel tests to an industry-relevant civil aircraft half-model designed by Airbus and in full-scale free-flight experiments with a Stemme S10 glider. The focus of this contribution is on closed-loop control. Moreover, it is shown that our approach can be used for very different setups. For this, the modeling of the dynamic responses of the two systems and the synthesis of robust controllers are compared. For the highly three-dimensional wind-tunnel model, differential pressures are used as surrogate control variables as a correlation with the lift can be described by a lookup table. Two control inputs, pulsed blowing, and two output variables are exploited. In the glider study, with a proFde that experiences a mostly two-dimensional flow, an approximate pressure gradient is used instead as a single output to avoid the necessity of such a lookup table. Active flow control is done again, exploiting pulsed blowing. For the controller synthesis, both in the single-input single-output and in the multivariable cases, the same approach is used, showing its versatility. In both setups, the closed-loop controllers succeed in adjusting the control inputs such that changing reference signals are followed, and the influence of disturbances are attenuated. For the civil aircraft model, completely new landing scenarios can be proposed using the closed-loop controller. [ABSTRACT FROM AUTHOR]

A low-speed wind-tunnel investigation has been undertaken to establish the effect of wing aspect ratio on Gurney-flap performance. Characterization is accomplished using a force balance and flow visualization. The Gurney-flap lift increment due to a shift in the zero-lift angle of attack was observed to scale with that of the lift-curve slope for different aspect ratios. As the aspect ratio reduced, a Gurney flap of greater height was required to maximize aerodynamic efficiency. The dependence of aerodynamic parameters (zero-lift angle of attack, minimum drag coefficient, and lift-curve slope) on the Gurney flap's height-to-chord ratio was also examined [ABSTRACT FROM AUTHOR]

The aerodynamic effects of the application of Gurney flaps of different heights and chordwise locations to five airfoils have been investigated in the Pennsylvania State University low-speed, low-turbulence wind tunnel. The effectiveness of each Gurney flap/airfoil combination is measured by the change in the maximum lift coefficient of the airfoil. When grouped by flap height and plotted against chordwise location, there is considerable scatter in the data, indicating that the effectiveness of the Gurney flap is strongly influenced by airfoil shape. Two anomalous cases are considered in detail. In the first case, the increase in cl, max is considerably different for two airfoils with the addition of a Gurney flap having the same height and mounted at the same chordwise location. The second case is one in which a Gurney flap of a specific height and mounting location is found to increase cl, maxs on one airfoil and decrease it on another. For these cases, pressure distributions, lift curves, and drag polars are provided. An explanation for the mechanism for the change in lift due to a Gurney flap is presented [ABSTRACT FROM AUTHOR]

The tests of the standard models serve to confirm the overall accuracy and repeatability of measurements in every single wind-tunnel facility and to confirm confidence in results obtained. The intention of the authors in this paper is to present, with exemplary test results, a methodology for certification and verification of the overall reliability of the T-38 Military Technical Institute's trisonic wind tunnel on the basis of check standard testing together with a small number of repeat-run sets prior to the customer's tests. Wind-tunnel data uncertainty is considered in the form of repeatability of a few presumably identical tests of a standard model. Test data are also correlated with those from other wind-tunnel facilities [ABSTRACT FROM AUTHOR]

Experimental studies were conducted by flapping a rigid rectangular wing with a mechanism that is capable of emulating complex insect wing kinematics, including figure-of-eight motions, in order to explore the fundamental unsteady flow on a flapping wing at micro-air-vehicle-scale Reynolds numbers. Force and moment measurements were obtained from a miniature six-component force transducer installed at the wing root. The wing was flapped in air and vacuum at the same frequency, and wing kinematics, and the resultant forces, were subtracted in order to obtain the pure aerodynamic forces. In the first part of this paper, the forces produced on the wing undergoing single-degree-of-freedom fixed-pitch pure flapping motions (no pitching or out-of-the-plane coning motions) were determined for a variety of pitch angles. The unsteady aerodynamic coefficients measured during these tests were almost six times the steady-state values measured in the wind tunnel. Flow visualization and particle image velocimetry tests were also conducted, which showed that the key reason for the force increase on the flapping wing is due to a strong leading-edge vortex for which the strength varied throughout the flapping cycle. In the second part of this paper, complete three-degree-of-freedom (flapping, pitching, and coning) insect wing kinematics were investigated for different pitching and coning variations. The aerodynamic forces obtained in these tests were compared with coefficients obtained from the single-degree-of-freedom flapping tests and wind-tunnel tests [ABSTRACT FROM AUTHOR]

The perforated test section walls of a transonic wind tunnel are responsible for generating edgetones, thus contributing significantly to the overall noise level and possibly invalidating the measurements of surface pressure fluctuations. In the present study, the wind tunnel was assembled with perforated top/bottom walls, solid side walls, and four perforated walls. The focus was to investigate the effects of tunnel background noise on compressible convex-corner flows, which correspond to the upper control surface of an aircraft wing. External acoustic disturbance has a minor effect on the mean surface pressure distributions, including flow expansion and recompression near the corner. Higher peak pressure fluctuations, which are associated with shock excursion phenomena, are observed with more intense tunnel background noise, particularly for the flow with initial separated boundary layers. [ABSTRACT FROM AUTHOR]

Turbulence grids are employed in an investigation to study the effects of freestream turbulence levels on aerodynamic loads for several low-aspect-ratio flat-plate wings. The objective of this research was to investigate how large levels of freestream turbulence affect a fixed wing in a low-speed turbulent flow. This paper specifically focuses on time-averaged mean load measurements of wings with three aspect ratios in highly turbulent and low-turbulence baseline flows. Relevant findings are that the presence of turbulence decreases the mean lift for an aspect ratio of 1 wing while increasing the mean lift for an aspect ratio of 2 and 4 wings. The stall angle of attack was increased for all aspect ratios investigated. Turbulence also modifies pitching moments to behave more smoothly than baseline cases. The standard deviations of the loads were found to be constant across the angle of attack, but the rolling moment was significantly increased with increasing aspect ratios. [ABSTRACT FROM AUTHOR]

The article discusses model-sidewall connection's impact for a static helicopter rotor blade airfoil experiment. It states that the gap between the wind-tunnel wall and airfoil can affect the three-dimensional near-wall flow topology for attack's static angle. It suggests the importance of understanding the model-sidewall interaction in improving wind-tunnel airfoil testing's accuracy and in reducing the sidewall interference.

This paper presents an improved structural analysis for a bearingless helicopter rotor. The bearingless rotor usually features a significantly large elastic twist in the flexbeam and additional unique structural characteristics. Thus, it will require sophisticated structural analysis and relevant numerical validation procedures due to its multiple load paths, as induced by the single or multiple fiexbeams and the torque tube. In this paper, an extended finite element formulation was derived to consider the multiple components as individual beam elements. A geometrically exact beam formulation was adopted to describe the nonlinear behavior of these major components in the rotor precisely. To implement the interconnecting kinematic relationship with the major components, Lagrange multipliers were used. The present static analysis was validated through comparisons with the existing multi-body dynamics analysis DYMORE. Additional results were obtained for rotating conditions in both a vacuum and a set atmosphere in a wind tunnel. Finally, an experimental full-scale bearingless rotor system was coupled with finite-state dynamic inflow aerodynamics to simulate forward flight. The present predictions of the rotor responses and aerodynamics were well correlated with those by CAMRAD II. [ABSTRACT FROM AUTHOR]

Wind tunnel measurements of the performance, airloads, and structural loads of a full-scale UH-60A Black Hawk main rotor operating at high advance ratios (up to 1.0) are compared with calculations obtained using the comprehensive rotorcraft analysis Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics II to understand physics and quantify this comprehensive code's accuracy and reliability in the prediction of rotor performance and loads at high-advance ratios. Detailed comparisons are made on rotor thrust, control angles, power, and section loads to illustrate and understand unique aeromechanics phenomena in this regime. The analysis correctly predicts the thrust reversal with collective at high advance ratios. Rotor induced plus profile power is also reasonably well predicted with proper modeling of the shank. Airloads and structural loads correlation is fair. A significant underprediction of 2-per-revolution structural loads is observed. [ABSTRACT FROM AUTHOR]

The article presents a study which examined the airplane-mode performances of the smart unmanned aerial vehicle (SUAV) using the nacelle-fixed auxiliary wing (FAWN). A wind-tunnel testing was performed to ascertain the aerodynamic configurations of the SUAV baseline and NFAW model. The endurance and range of wing body and specific fuel consumption (SFC) of the SUAV are determined.

The article presents a study which examined the lift of a biplane based on computations of the gap g, the vertical distance between the quarter chords of the two wings perpendicular to the free stream, the stagger distances, aspect ratio and angle of attack. The wind-tunnel at the University of Dayton was used, with Reynolds number as the basis of these computations. Equations for linear regression on the stagger-dependent lift-curve slopes and y-intercepts are offered.

This contribution discusses the implementation of active flow-separation control for a three-dimensional high-lift wing-body configuration under atmospheric low-speed wind-tunnel conditions. The slot actuators are applied on the suction side of the trailing-edge flap to prevent local flow separation. The experimental results indicate that the pulsed blowing flow control technique is effective on the present configuration with a global performance enhancement. Numerical investigations are the focus of this article. The baseline case is characterized by substantial portions of separated flow. Thus, the influence of grid resolution and turbulence modeling is investigated. Based on this an intermediate mesh in combination with the Shear Stress Transport model gives the best compromise between quality and computational turnaround times. The steady Reynolds Averaged Navier Stokes (RANS) calculations carried out with constant blowing demonstrate the feasibility to simulate active flow control concepts. The key flow control method is the pulsed blowing. The verification of the unsteady RANS approach with active flow control shows that high computational resources are required for consistent numerical evaluations. The computational results highlight the ability of pulsed blowing at moderate blowing momentum coefficients to suppress the flow separation on the traillng-edge flap. The numerical results show an acceptable agreement with the experiments. [ABSTRACT FROM AUTHOR]

In the 1970s the B-747-100 alleviation tests, sponsored by NASA with FAA participation, provided unique continuous-wave lidar data, in which the vortex cores could be measured because they were marked with smoke. Measurements for both normal and alleviated wakes showed that the vortex core is persistent whereas the outer circulation of the vortex decays. This form of decay is exactly the opposite of what is observed at low Reynolds number for small aircraft or in wind tunnels. The stability of the core is consistent with the essentially laminar core observed via flow visualization. The circulation profile for the normal wake is well fitted by the Burnham-Hallock model both after initial rollup and after significant decay of the total circulation. The core size for the normal B-747-100 wake is larger (5 % of wingspan) than typically observed for aircraft with zero or two wing-mounted engines. The observed wake decay means that the total vortex circulation is approximately equal to the popular vortex strength parameter: average circulation over vortex radii from 5 to 15 m. [ABSTRACT FROM AUTHOR]

A half model of a scaled-down aircraft is designed and tested in the wind tunnel to validate the mathematical model of uncertainty for unsteady pressure coefficients in the frequency domain. In the wind-tunnel test a step-swept test was conducted to obtain the frequency response function. Then a time-domain response test was performed with turbulence excitation to identify the aircraft's on-line poles. Based on the tested frequency response function for the on-line poles the structured singular value (µ) method was applied to determine the aerodynamic uncertainty level. Finally, the widely used µ analysis method was employed to compute the worst-case flutter boundary, and the result was compared with the experimental flutter velocity. The experimental flutter velocity (30.5 m/s) is in the range of the predicted robust flutter boundary (28.5 m/s) in which parameter uncertainties were taken into account in the numerical model. Experimental results validate that the uncertainty quantification theoretical frameworks incorporating experimental data can estimate the proper aerodynamic uncertainty level and predict a safe flutter boundary. The present results suggest that the time-response validation theoretical framework is more advantageous in robust stability analysis than the one based upon the frequency response function. [ABSTRACT FROM AUTHOR]

The flow at high angles of attack over a slender body with a short nose is studied. The effects of nose geometry on the flow and its control using a strake are investigated. Different nose geometries are individually connected to a common cylindrical aftbody to result in three models with an overall length-to-diameter ratio of 8.4. Wind-tunnel experiments are conducted at a freestream velocity of 18.5 m/s and angles of attack ranging from -40 to 60 deg to investigate both time-averaged and unsteady behaviors. The results are compared with those from a previous study on models with two of the same nose geometries but overall length-to-diameter ratio of 4. Additional to the commonly observed average vortex asymmetry, an oscillation of the aftbody vortices is observed even at moderate angles of attack. This oscillation is suppressed by just a slight "pointing" of the nose. Additionally a seemingly blunt nose can actually behave like a pointed nose. [ABSTRACT FROM AUTHOR]

The article discusses a study which demonstrated the usefulness of a test technique for the measurement of intake drag of air-breathing vehicles in a limited Mach number range. It illustrates the model mounted in the base flow wind tunnel. It describes the technique as consisting of a metric intake body and nonmetric portion of the intake model to the annular strain gauge balance. The study observed that the technique can be used to evaluate and optimize the performance of the intake system.

A low-speed wind-tunnel investigation is presented, exploring the effect of delta wings configured in an 'X' biplane geometry. Wing leading-edge sweeps ranging from 50 to 70 deg and anhedral/dihedral angles up to 40 deg were explored. Results indicated a significant attenuation of vortex lift production, while the attached flow lift component was enhanced due to nonplanarity, an outcome supported by theory developed to analytically simulate this wing geometry. Observed benefits of this wing configuration amplify with wing slenderness, where marked increases in the range parameter may be realized. Both theory and experiment indicate that the optimal anhedralldihedral angles for the tested geometries ranged from 20 to 30 deg. [ABSTRACT FROM AUTHOR]

This paper deals with the application of a via-swashplate higher harmonic control for the reduction of vibratory loads transmitted to tiltrotor fuselage by the proprotor-pylon-wing system. The control laws are derived through a quadratic optimal control methodology, based on the predictions given by an aeroelastic simulation tool developed in the past by the authors. It uses nonlinear beam theory to model the structural behavior of proprotor blades and wing, and a potential aerodynamics boundary-element approach suitable to capture body-vortex interactions effects to evaluate wake inflow. Considering a XV-15-1ike configuration in airplane- and helicopter-mode flight, the numerical investigation examines different strategies in the identification of the optimal control law, analyzing their effectiveness and robustness in reducing the wing-root vibratory loads. [ABSTRACT FROM AUTHOR]

The steep-approach method is identified to be an operational measure for transport aircraft to reduce perceived noise. The steep-approach performance of multifunctional aileron devices compared to a conventional spoiler configuration was investigated, and the feasibility of the concepts to maintain roll control of a conventional aileron was evaluated. Multifunctionai aileron devices comprise a horizontally segmented aileron, designated as a deceleron, and a laterally segmented aileron, designated as a split aileron. These aileron devices cover the same dimensions as a conventional aileron. A detailed wind-tunnel half-model in a high-lift configuration was used and the aerodynamic coefficients were obtained for various angles of attack. The tests were conducted in a low-speed wind-tunnel facilitywith a Reynolds number of 106 Aileron deflection and differential aileron deflection angles were varied. The most relevant parameters were the minimum approach velocity, maximum descent velocity, and maximum approach angle. Regarding the objective to maximize the approach angle, it was found that none of the multifunctional ailerons were superior to conventional spoilers. However, the multifunctional ailerons do not have the penalty of a higher approach speed for steep approach, being superior in this manner to the conventional spoiler. The roll control ability was maintained by the new ailerons for moderate differential deflections. [ABSTRACT FROM AUTHOR]

A series of three wind-tunnel static and forced-oscillation tests were conducted on a model of a generic unmanned combat air vehicle. These tests are part of an international research effort to assess the state of the art of computational fluid dynamics methods to predict the static and dynamic stability and control characteristics. The experimental data set includes not only force and moment time histories, but also surface pressure and offbody particle image velocimetry measurements. The extent of the data precludes a full examination within the scope of this paper. This paper provides a general description and selected examples of the available static and dynamic data, as well as some of the observed trends. [ABSTRACT FROM AUTHOR]

A comprehensive research program designed to investigate the ability of computational methods to predict stability and control characteristics of a generic unmanned combat air vehicle has been undertaken. The integrated approach to simulating static and dynamic stability characteristics was performed by the NATO Research and Technology Organization Task Group AVT-161. The vehicle named Stability and Control CONfiguration (SACCON) was the subject of an intensive computational and experimental study. The stability characteristics of the vehicle were evaluated via a highly integrated approach, where computational fluid dynamics and experimental results were used in a parallel and collaborative fashion. The results show that computational methods have made great strides in predicting static and dynamic stability characteristics, but several key issues need to be resolved before efficient, affordable, and reliable predictions are available. [ABSTRACT FROM AUTHOR]

This paper considers the ways in which structural model parameter variability can influence aeroelastic stability. Previous work on formulating the stability calculation (with the Euler equations providing the aerodynamic predictions) is exploited to use Monte Carlo, interval, and perturbation calculations to allow this question to be investigated. Three routes are idenitfied. The first involves variable normal-mode frequencies only. The second involves normal-mode frequencies and shapes. Finally, the tbhird, in addition to normal-mode frequencies and shapes, also includes their influence on the static equilibrium. Previous work has suggested only considering the first route, which allows significant gains in computational efficiency if reduced-order models can be built for the aerodynamics. However, results in the current paper show that neglecting the mode-shape variation can give misleading results for the flutter-onset prediction, complicating the development of reduced aerodynamic models for variability analysis. [ABSTRACT FROM AUTHOR]

A method is proposed to estimate the lift curves or changes of lift curves of a circulation-control wing at low speeds such as the ones encountered during takeoff and approach. This method is for rapid prototyping during the conceptual design stage with a goal to predict the performance or the change of lift coefficient within 5-10% of accuracy. This method includes three parts: an alternate approach for estimating jet speed without assuming the jet static pressure at the exit, an empirical approach for estimating the lift coefficients of circulation-control airfoils, and a lifting-fine theory for estimating wing-lift curves based on the airfoil-lift data. The validation against the wind-tunnel test data available for a modified Intruder A-6 aircraft with a circulation-control wing was deemed satisfactory. [ABSTRACT FROM AUTHOR]

The process of performing in-flight measurements of the canard wake on a Velocity XL-5 retractable gear (RG) is detailed and the results are compared to computational models of the flowfield. A combination of an 18-probe total pressure rake and two five-hole probes were used. Preflight wind-tunnel tests were used for probe calibration and a three-dimensional viscous code was used for the flowfield computations. The comparisons between the computations and the measurements were very good. The resulting database validated the computations and provides a reference for future flight experiments. [ABSTRACT FROM AUTHOR]

The effects of unsteady aerodynamics for a tandem-wing unmanned aerial vehicle (UAV) prototype during the morphing stage have been investigated numerically at low Reynolds numbers (Re < 106). The morphing stage consists of two stages. The first stage involves the unfolding of the canards and wings from their initial position in which they were attached and aligned parallel to the UAV fuselage. The second stage involves the rotation of the unfolded canards and wings around the chord axis toward their final orientations. Lift and drag coefficients for the canard and wing vary significantly, particularly during the second morphing stage. Maximum deviation can be 2 to 14 times higher than the final steady state values. A significant increase in the hinge moment is also observed during the first morphing stage but the sign of the moment remains the same throughout. The unsteady aerodynamics, however, have larger effects during the second morphing stage, including the rapid change in the sign for the hinge moment, the corresponding magnitude also differs by (+) 57% and (+) 260% for the canard and wing, respectively. Attempts have been made to correlate these observations with the vortex interactions between the canard and wing during the morphing process. [ABSTRACT FROM AUTHOR]

Ideal flight sheds the least amount of kinetic energy into a wake while imparting momentum sufficient to balance the vehicle weight. This combination defines a unique downwash distribution for the wake, which an aircraft designer should provide for. A central fuselage, as required for the typical flight objective, presents an obstacle to this intent. A wing interrupted by a prominent fuselage is expected to shed inboard trailing vortices with central upwash harmful to the span efficiency of the aircraft. It is proposed here that a trailing edge on the fuselage can be used to control the circulation in the central region of the aircraft so that the central downwash deficiency can be avoided. Such a Kutta edge can further be applied as part of a high-lift system to increase central downwash by increasing the loading on the wing root and lift over the fuselage itself. Time-averaged flowfields behind a wing-body combination with and without a Kutta edge have been measured in wind-tunnel experiments. The results show that an edged aft-body does influence central circulation, as predicted. Flight with ideal wakes may be more readily attained than hitherto realized. [ABSTRACT FROM AUTHOR]

This paper describes a fundamental experimental study of the flow structure around a single three-dimensional (3D) transonic shock control bump (SCB) mounted on a flat surface in a wind tunnel. Tests have been carried out with a Mach 1.3 normal shock wave located at a number of streamwise positions relative to the SCB. Details of the flow have been studied using the experimental techniques of schlieren photography, surface oil flow visualization, pressure sensitive paint, and laser Doppler anemometry. The results of the work build on the findings of previous researchers and shed new light on the flow physics of 3D SCBs. It is found that spanwise pressure gradients across the SCB ramp and the shape of the SCB sides affect the magnitude and uniformity of flow turning generated by the bump, which can impact on the spanwise propagation of the quasi-two-dimensional (2D) shock structure produced by a 3D SCB. At the bump crest, vortices can form if the pressure on the crest is significantly lower than at either side of the bump. The trajectories of these vortices, which are relatively weak, are strongly influenced by any spanwise pressure gradients across the bump tail. A significant difference between 2D and 3D SCBs highlighted by the study is the impact of spanwise pressure gradients on 3D SCB performance. The magnitude of these spanwise pressure gradients is determined largely by SCB geometry and shock position. [ABSTRACT FROM AUTHOR]

Design informatics has been proposed as a method of achieving efficient design. Design informatics allows efficient exploration of the design space and implements conception support for innovative design concepts. In the present study, this approach was applied to the design of the full configuration of the silent supersonic technology demonstrator. The design-informatics approach involves two steps: optimization and data mining. As a first step, multidisciplinary design optimization with multiple objectives was performed considering aerodynamics, stability, structures, aeroelasticity, and boom noise using an evolutionary algorithm. As a second step, data mining was carried out using a self-organizing map and an analysis of variance to extract design knowledge from the acquired optimization results for determining a conclusive compromise solution. Although the optimization had insufficient convergence due to the constraints of time and computational cost, the data mining brought important design information. Since the design space was consequently clarified, a compromise solution was successfully determined. Thus, design informatics is essential to an efficient design process. [ABSTRACT FROM AUTHOR]

The aerodynamics of parachute systems are highly unsteady, resulting from separated flows, vortex shedding, etc., caused by complex interactions between a flexible, bluff canopy and its payload. Nevertheless, and when dynamically permitted, the flowfields and surface forces have often been approximated using the assumptions and characteristics of rigid bodies and steady-state flows. The validity of the steady-state assumption is considered here in the case of the seldom-studied annular parachute in constant-speed descent. Computational fluid dynamics simulations obtained by solving the Reynolds-averaged Navier-Stokes equations in a steady-state manner are compared with those calculated in a transient manner. The simulations were performed at Reynolds numbers of 8.03 · 106 and 10.04 · 106, around two rigid and vertically offset concentric rings of unequal diameters. Comparisons of the pressure, velocity, vorticity, and turbulence contours show many important features of the transient simulations not being captured in the steady-state calculations. Those include Kelvin-Helmhoitz instabilities near the edges of the two annuli and their interaction with the boundary-layer vorticity and the jet emerging from the interannulus gap. This suggests that steady-state simulations around annular parachutes may only be appropriate for total drag calculations, and only if the flow velocity is high enough. [ABSTRACT FROM AUTHOR]