Showing papers in "Journal of Aircraft in 2005"

Efficient Optimization Design Method Using Kriging Model

Abstract: The Kriging-based genetic algorithm is applied to aerodynamic design problems. The Kriging model, one of the response surface models, represents a relationship between the objective function (output) and design variables (input) using stochastic process. The kriging model drastically reduces the computational time required for objective function evaluation in the optimization (optimum searching) process. ‘Expected improvement (EI)’ is used as a criterion to select additional sample points. This makes it possible not only to improve the accuracy of the response surface but also to explore the global optimum efficiently. The functional analysis of variance (ANOVA) is conducted to evaluate the influence of each design variable and their interactions to the objective function. Based on the result of the functional ANOVA, designers can reduce the number of design variables by eliminating those that have small effect on the objective function. In this paper, the present method is applied to a two-dimensional airfoil design and the prediction of flap’s position in a multi-element airfoil, where the lift-to-drag ratio (L/D) is maximized.

Nonlinear Flight Dynamics of Very Flexible Aircraft

Abstract: This paper focuses on the characterization of the response of a very ∞exible aircraft in ∞ight. The 6-DOF equations of motion of a reference point on the aircraft are coupled with the aeroelastic equations that govern the geometrically nonlinear structural response of the vehicle. A low-order strain-based nonlinear structural analysis coupled with unsteady flnite-state potential ∞ow aerodynamics form the basis for the aeroelastic model. The nonlinear beam structural model assumes constant strain over an element in extension, twist, and in/out of plane bending. The geometrically nonlinear structural formulation, the flnite state aerodynamic model, and the nonlinear rigid body equations together provide a low-order complete nonlinear aircraft analysis tool. The equations of motion are integrated using an implicit modifled generalized-alpha method. The method incorporates both flrst and second order nonlinear equations without the necessity of transforming the equations to flrst order and incorporates a Newton-Raphson sub-iteration scheme at each time step. Using the developed tool, analyses and simulations can be conducted which encompass nonlinear rigid body, nonlinear rigid body coupled with linearized structural solutions, and full nonlinear rigid body and structural solutions. Simulations are presented which highlight the importance of nonlinear structural modeling as compared to rigid body and linearized structural analyses in a representative High Altitude Long Endurance (HALE) vehicle. Results show signiflcant difierences in the three reference point axes (pitch, roll, and yaw) not previously captured by linearized or rigid body approaches. The simulations using both full and empty fuel states include level gliding descent, low-pass flltered square aileron input rolling/gliding descent, and low-pass square elevator input gliding descent. Results are compared for rigid body, linearized structural, and nonlinear structural response.

Review of Unsteady Vortex Flows over Slender Delta Wings

Abstract: Nomenclature AR = aspect ratio; amplitude ratio B = probability density function of velocity c = root chord length f = frequency k = reduced frequency; axial wave number n =w ave number in angular direction P = probability p = pressure fluctuation Re =R eynolds number based on chord length S = spectral density s = local semispan T = period t = time U∞ = freestream velocity u = axial velocity v = swirl velocity x = streamwise distance xbd = breakdown location y = spanwise distance z =v ertical distance above wing surface α = angle of attack � = circulation δ = flap angle � = sweep angle ν = kinematic viscosity τ = time constant � =fi nangle ω =v orticity; radial frequency

Aerodynamic and Static Aeroelastic Characteristics of a Variable-Span Morphing Wing

Abstract: The morphing concept for unmanned aerial vehicles is a topic of current research interest in aerospace engineering. One concept of morphing is to change the wing configuration during flight to allow for multiple flight regimes. A particular approach to planform morphing is a variable-span morphing wing to increase wingspan to reduce induced drag and increase range and endurance. The wing area and the aspect ratio of the variable-span morphing wing increase as the wingspan increases. This means that the total lift increases while the induced drag is reduced, whereas the wing-root bending moment increases, thus, requiring a larger bending stiffness of the wing structure. Therefore, a study of the variable-span morphing wing requires not only aerodynamic analysis but also an investigation of the aeroelastic characteristics of the wing. The aerodynamic characteristics of the variable-span morphing wing are investigated, and a static aeroelastic analysis is performed.

Flight Characteristics of Shaping the Membrane Wing of a Micro Air Vehicle

Abstract: Biologically inspired concepts are rapidly expanding the range of aircraft technology. Consideration is given to merging two biologically-inspired concepts, morphing and micro air vehicles, and the resulting flight characteristics are investigated. Specifically, wing shaping is used to morph the membrane wings of a micro air vehicle. The micro air vehicle has poor lateral control because hinges, and consequently ailerons, are difficult to install on a membrane wing. Instead, a set of torque rods, aligned along the wings, are used to twist the membrane and shape the wing. The resulting morphing is shown to provide significant control authority for lateral dynamics. A set of flight tests are undertaken to determine the flight characteristics by commanding pulses and doublets to the control actuation. The vehicle demonstrates excellent roll performance in response to wing shaping. Futhermore, the vehicle demonstrates several types of spin behavior related to combinations of elevator deflection and the wing shaping.

Vortex-Generator Model and Its Application to Flow Control

Abstract: A new vortex-generator model is introduced, the jBAY model, which provides an efficient method for computational-fluid-dynamics (CFD) simulation of flow systems with vortex-generator arrays. The ...

Application of System Identification to Aircraft at NASA Langley Research Center

Abstract: The past, present, and future of system identification applied to aircraft at NASA Langley Research Center (LaRC) in Hampton, Virginia, are discussed. Significant research advances generated at NASA LaRC in the past are summarized, including some perspective on the role these developments played in the practice of system identification applied to aircraft. Selected recent research efforts are described, to give an idea of the type of activities currently being pursued at NASA LaRC. These efforts include real-time parameter estimation, identifying flying qualities models, advanced experiment design and modeling techniques for static wind-tunnel database development, and indicial function identification for unsteady aerodynamic modeling. Projected future developments in the area are outlined

Aircraft Structural Morphing Using Tendon-Actuated Compliant Cellular Trusses

Abstract: Recently, smoothly-deforming aircraft structures have been investigated for their ability to adapt to varying flight conditions. Researchers aim to achieve large changes in the shape of the wings: area changes of up to 50% and aspect-ratio changes of up to 200% are being pursued. The research described in this paper aims to develop a structural concept capable of achieving continuous stable deformations over a large range of aircraft shapes. The basic concept underlying the approach is a compliant cellular truss, with tendons used as active elements. The truss members of the unit cell are connected through compliant joints such that only modest bending moments may be transmitted from one member to another. Actuation is achieved by pulling on one set of cables while releasing another set. The tendonactuated compliant truss can be made to behave locally, and temporarily, as a nearmechanism, by releasing appropriate cables. As a result, in the absence of aerodynamic forces, the structure can be morphed using relatively low forces. The cables are reeled in or released in a controlled manner while the structure is loaded, hence, the stability of the structure can be maintained in any intermediate position. Highly-distributed actuation also enables the simultaneous achievement of global shape changes as the accumulation of local ones, while the use of compliant joints rather than true rotating joints eliminates binding as a significant concern. A six-noded octahedral cell with diagonal tendon actuation is developed for a bending type deformation in the wing. Initial cell geometry is determined by “strain matching” to the local morphing deformation required by the application. A finite element analysis is performed on a wing made of these unit cells and sized for a representative UAV weighing 3000 lbs. The areas of the individual truss members are sized so that they don’t fail or buckle under the air loads, while deflection at the wing tip is reduced. The octahedral unit cell is capable of achieving smooth deformations of the truss structure. The cell size is dictated by the available space and the morphing strain. The cell sizes are reasonable for strains on the order of 10% to 15% and get smaller for larger strains. Additional cell shapes are being investigated for larger area changes through a process of topology optimization using genetic algorithms. Numerous other technical challenges remain, including the details of actuation and a robust skin.

Semi-Active Vibration Isolation Using Magnetorheological Isolators

Abstract: An experimental and theoretical analysis of a vibration isolation system that uses magnetorheological (MR) fluid-based semi-active isolators is presented. In doing so, a vibration isolator that uses MR fluids is designed, manufactured, and experimentally evaluated. Typically, a Bingham-plastic model, with its accompanying zero speed force step discontinuity, would be used to model an MR isolator. A new nonlinear model, with simple low-speed hysteresis characteristics, is proposed to describe the hysteresis force characteristics of the MR isolator. The damping forces of the MR isolator with different excitation frequency and current input are measured and compared with that resulting from the hysteresis model for the verification of the theoretical analysis. A vibration isolation system with the MR isolator is constructed, and its dynamic equation of motion is derived. A simple skyhook controller is formulated to attenuate the vibration of the system. Controlled performances of the vibration isolation system are experimentally and theoretically evaluated in the frequency and time domains. A key conclusion is that a simple model of the low-speed force vs low-speed velocity hysteresis characteristics is necessary for successful prediction of open- and closed-loop performance in both the time and frequency domains.

Dynamic Response of Aeroservoelastic Systems to Gust Excitation

Abstract: Frequency-domain and time-domain approaches to dynamic response analysis of aeroservoelastic systems to atmospheric gust excitations are presented. The discrete and continuous gust inputs are defined in either time-domain or stochastic terms. The various options are formulated in a way that accommodates linear control systems of the most general form. The frequency-domain approach is based on the interpolation of generalized aerodynamic force coefficient matrices and the application of Fourier transforms for time-domain solutions. The time-domain approach uses state-space formulation that requires the frequency-dependent aerodynamic coefficients to be approximated by rational functions of the Laplace variable. Once constructed, the state-space equations of motion are more suitable for time simulations and for the interaction with control design algorithms. However, there is some accuracy loss because of the rational approximation. The spiral nature in the complex plane of the gust-related aerodynamic terms is discussed, and means are provided for dealing with the associated numerical difficulties. A hybrid formulation that does not require the rational approximation of the gust coefficients is also presented for optional use in discrete gust response analysis. The various methods were utilized in the ZAERO software and applied to a generic transport aircraft model.

Joined-Wing Aeroelastic Design with Geometric Nonlinearity

Abstract: An integrated process is presented that advances the design of an aeroelastic joined-wing concept by incorporating physics-based results at the system level. For instance, this process replaces empirical mass estimation with high-fidelity analytical mass estimations. Elements of nonlinear structures, aerodynamics, and aeroelastic analyses were incorporated with vehicle configuration design. This process represents a significantly complex application of aeroelastic structural optimization. Specific fuel consumption for a fixed lift-to-drag ratio was considered in the process for estimating fuel to size the structure to meet range and loiter requirements. This design process was implemented on a single configuration for which two crucial nonlinear phenomena contribute to structural failure: large deformation aerodynamics and geometrically nonlinear structures. A correct model of the nonlinear aeroelastic physics offers the possibility of a successful design. Unconventional features of a joined-wing concept are presented with the aid of this unique design model. Hopefully, insight derived from the nonlinear aeroelastic design might be leveraged to the benefit of future joined-wing designs.

Delta wing stall and roll control using segmented piezoelectric fluidic actuators

Abstract: The separated flow around a balance-mounted, 60-deg sweptback, semispan delta wing with a sharp leading edge was controlled using zero-mass-flux periodic excitation from a segmented leading-edge slot. Excitation was generated by cavity-installed piezoelectric actuators operating at resonance with amplitude modulation (AM) and burst mode (BM) signals being used to achieve reduced frequencies (scaled with the freestream velocity and the root chord) in the range from O(1) to O(10). Results of a parametric investigation, studying the effects of AM frequency, BM duty cycle and frequency, excitation amplitude, location of the actuation along the leading edge, and optimal phase difference between the actuators, as well as the Reynolds number, are reported and discussed

Aeroelastic System Development Using Proper Orthogonal Decomposition and Volterra Theory

Abstract: This research combines Volterra theory and proper orthogonal decomposition (POD) into a hybrid methodology for reduced-order modeling of aeroelastic systems. The outcome of the method is a set of linear ordinary difierential equations (ODEs) describing the modal amplitudes associated with both the structural modes and the POD basis functions for the ∞uid. For this research, the structural modes are sine waves of varying frequency, and the Volterra-POD approach is applied to the ∞uid dynamics equations. The structural modes are treated as forcing terms which are impulsed as part of the ∞uid model realization. Using this approach, structural and ∞uid operators are coupled into a single aeroelastic operator. This coupling converts a free boundary ∞uid problem into an initial value problem, while preserving the parameter (or parameters) of interest for sensitivity analysis. The approach is applied to an elastic panel in supersonic cross ∞ow. The hybrid Volterra-POD approach provides a low-order ∞uid model in state-space form. The linear ∞uid model is tightly coupled with a nonlinear panel model using an implicit integration scheme. The resulting aeroelastic model provides correct limit-cycle oscillation prediction over a wide range of panel dynamic pressure values. Time integration of the reduced-order aeroelastic model is four orders of magnitude faster than the high-order solution procedure developed for this research using traditional ∞uid and structural solvers.

Biodynamic response mitigation to shock loads using magnetorheological helicopter crew seat suspensions

Abstract: Biodynamic response mitigation of both sinusoidal vibration and shock loads using a magnetorheological (MR) seat suspension is investigated. In doing so, an MR seat suspension model for helicopters, with a detailed lumped parameter model of a human body, was developed. The lumped parameter model of the human body consists of four parts: pelvis, upper torso, viscera, and head. From the model, the governing equation of motion of the MR seat suspension considering the human body was derived. Based on this equation, a semi-active nonlinear optimal control algorithm appropriate for the MR seat suspension was developed. The simulated control performance of the MR seat suspension was evaluated under both sinusoidal vibration and shock loads due to a vertical crash landing of a helicopter. In addition, the mitigation of injuries to humans due to such shock loads was also evaluated and compared with that of the passive seat suspension using a passive hydraulic damper.

Spanwise Lift Distribution for Blended Wing Body Aircraft.

Abstract: A study is presented of the effects of spanwise lift distribution on aerodynamic efficiency for a blended wing body (BWB) configuration of a given baseline planform. The baseline geometry is initially assessed by a high-fidelity aerodynamic model based on a multiblock structured grid Reynolds-averaged Navier-Stokes (RANS) solution. The accuracy of the simulation is investigated by a grid sensitivity study regarding total drag and its pressure drag and skin-friction drag components. Excessive outer wing loading with associated shock wave, hence, wave drag, has been revealed to be the major factor degrading the aerodynamic performance of the baseline BWB model. To relieve the outer wing, an efficient low-fidelity panel method aerodynamic model is used for the inverse design to achieve the target lift distributions through the variation of twist distribution along the span, shifting the load inboard. For the given BWB geometry, the baseline model is retwisted to achieve an elliptic, a triangular, and an averaged elliptic/triangular spanwise loading distribution. The designs are then analyzed using the high-fidelity RANS aerodynamic model. The wave drag component of the total drag for different span loadings is extracted from the flowfield solution to gain insight into the drag reduction provided by the new twist designs

Flow Visualization of Micro Air Vehicle Scaled Insect-Based Flapping Wings.

Abstract: Afl ow-visualization experiment was conducted on an insect-based flapping-wing mechanism. This enabled greater understanding and insight to be gained on the unsteady aerodynamic phenomena that are responsible for the enhanced lift of wings operating at low Reynolds numbers in hovering flapping flight. Flow-visualization images were acquired with a strobbed laser sheet to illuminate the flow, which was seeded with a mineral oil fog. The general flowfield structure was found to consist of a folded wake, with a relatively large starting vortex at the end of each half-stroke. A large flow recirculation region was generated in the plane of flapping, which was centered around the two extreme flapping displacements. These general flowfield features were enhanced by detailed observations of the local flowfield around the wing section. One observation was the presence of multiple vortices on the top surface of the wing as it underwent translation. Furthermore, the local flowfield images clearly showed the growth of the leading-edge vortex as a function of span and identified the presence of separated flow on the outboard regions of the wing. These experimental results were supported by a free vortex modeling of the wake developments. The model was found to predict similar wake flowfield dynamics to that found in the experiments. This research has contributed to a better understanding of the unsteady aerodynamic mechanisms that are responsible for the enhanced lift of insect-based flapping wings in hover.

Numerical Investigation of Low-Aspect-Ratio Wings at Low Reynolds Numbers

Abstract: A numerical analysis is performed to study the flow around low-aspect-ratio (LAR) wings and more particularly the resulting lift-and-drag force. The research is focused on low-Reynolds-number aerodynamics, as LAR wings are crucial for the development of microair vehicles (MAVs). The flow around LAR wings is characterized by complex three-dimensional flow phenomena. These phenomena include wing-tip vortices, flow separation and reattachment, laminar to turbulent transition, and a mutual interaction among these phenomena. The flow is studied using a commercial computational fluid dynamics (CFD) program and a strip method. The CFD code is used to investigate the three-dimensional flow aerodynamics of rectangular LAR wings with an aspect ratio between 0.5 and 2 at a Reynolds number of 1 x 10 5 . Simulations on a flat plate and a reflex-type low-Reynolds-number profile (S5010), which is representative for a flying-wing MAV, are performed and compared. Experimental data is used for comparison and validation. The effects of flow separation and low Reynolds numbers are further investigated using a strip method. Two accurate formalized methods to predict lift and drag are derived. The first method is applicable to profiled wings with moderate low-Reynolds-number effects. The second method, which is based on the strip method, is more general and is also applicable to flat plates and wings exhibiting large regions of flow separation.

Simulation of Helicopter Shipboard Launch and Recovery with Time-Accurate Airwakes

Abstract: A simulation of the helicopter/ship dynamic interface has been developed and applied to simulate a UH-60A operating from an LHA class ship. Time accurate CFD solutions of the LHA airwake are interfaced with a flight dynamics simulation based on the GENHEL model. The flight dynamics model was updated to include improved inflow modeling and gust penetration effects of the ship airwake. A maneuver controller was used to simulate pilot control inputs for specified approach and departure trajectories. The CFD solutions show significant time varying flow effects in the airwake. Time histories of the aircraft angular rate and pilot control activity indicate that the time varying nature of the airwake has significant effect on aircraft response and pilot workload.

Aircraft Spin Recovery, with and without Thrust Vectoring, Using Nonlinear Dynamic Inversion

Abstract: The present paper addresses the problem of spin recovery of an aircraft as a nonlinear inverse dynamics problem of determining the control inputs that need to be applied to transfer the aircraft from a spin state to a level trim flight condition. A stable, oscillatory, flat, left spin state is first identified from a standard bifurcation analysis of the aircraft model considered, and this is chosen as the starting point for all recovery attempts. Three different symmetric, level-flight trim states, representative of high, moderate, and low-angle-of-attack trims for the chosen aircraft model, are computed by using an extended-bifurcation-analysis procedure. A standard form of the nonlinear dynamic inversion algorithm is implemented to recover the aircraft from the oscillatory spin state to each of the selected level trims. The required control inputs in each case, obtained by solving the inverse problem, are compared against each other and with the standard recovery procedure for a modern, low-aspect-ratio, fuselage heavy configuration. The spin recovery procedure is seen to be restricted because of limitations in control surface deflections and rates and because of loss of control effectiveness at high angles of attack. In particular, these restrictions adversely affect attempts at recovery directly from high-angle-of-attack oscillatory spins to low-angleof-attack trims using only aerodynamic controls. Further, two different control strategies are examined in an effort to overcome difficulties in spin recovery because of these restrictions. The first strategy uses an indirect, two-step recovery procedure in which the airplane is first recovered to a high- or moderate-angle-of-attack level-flight trim condition, followed by a second step where the airplane is then transitioned to the desired low-angle-of-attack trim. The second strategy involves the use of thrust-vectoring controls in addition to the standard aerodynamic control surfaces to directly recover the aircraft from high-angle-of-attack oscillatory spin to a low-angle-of-attack level-flight trim state. Our studies reveal that both strategies are successful, highlighting the importance of effective thrust management in conjunction with suitable use of all of the aerodynamic control surfaces for spin recovery strategies.

Case for Small Supersonic Civil Aircraft

Abstract: Civil aviation progress in the last 40 years has included a significant expansion of the small civil aircraft market involving regional jets, business jets, and the emerging personal jets. A significant factor in the growth of the small civil aircraft market is the value of time. Recognition of the ever-increasing value of time has lead to increased interest in the feasibility of a small supersonic civil aircraft. The step to supersonic speeds offers the potential of a dramatic decrease in travel time. Feasibility studies of a small quiet supersonic jet (QSJ) have been conducted. Market research, environmental concerns, program and design requirements, and vehicle characteristics are summarized. Areas for concentrated future supersonic aeronautics research and development efforts are highlighted.

Effect of Atmospheric Transmission and Radiance on Aircraft Infrared Signatures

Abstract: This paper elaborates the role of atmosphere in determining infrared signatures of aircraft as perceived by a ground-based infrared detector, similar to the case of an infrared guided heat-seeking surface-to-air missile. The main objectives are to assess the effect of atmosphere on aircraft infrared signature and to evaluate infrared bands in which a conventional fighter-class aircraft is most susceptible to infrared guided missiles. Such an analysis is of paramount importance for aircraft infrared signature management and aircraft mission planning. First, the lock-on range is derived as a function of aircraft, missile seeker, and atmospheric parameters. Thus, the role of atmospheric radiance and transmittance in determining aircraft signature level as perceived by the missile infrared seeker is brought out. The role of various atmospheric constituents in dictating infrared characteristics of the atmosphere is also discussed. The Berger’s model is used for computing atmospheric/sky radiance and the Lowtran-7 model to compute atmospheric transmissivity. The infrared bands in which the aircraft signature is prominent are identified, and the variation of aircraft signature level and lock-on range with respect to a typical surface-to-air missile within these bands are analyzed and discussed for a representative case.

In-flight wing deformation characteristics during limit-cycle oscillations

Abstract: Oscillatory wing response data were measured on an F-16C aircraft during limit-cycle-oscillation (LCO) testing of an external store configuration. The configuration tested exhibited typical LCO response in the transonic flight regime. Deformation characteristics were measured at 11 locations on the wing and missile launchers during various LCO events. These measurements allowed viewing of the aeroelastic mode of instability for various flight conditions. Details of a linear flutter analysis model are presented, and the predicted eigenmode from the linear flutter analysis is compared to the in-flight measured mode. The measured LCO modes are also compared for level flight vs elevated load factor flight and subcritical LCO vs critical LCO conditions. At the onset of LCO, the mode shape bears a strong resemblance to the flutter mode predicted by linear flutter analysis. Further, the wing deformation characteristics during LCO vary significantly with respect to Mach number and load factor. The nonsynchronous motion of the LCO diminishes and becomes more synchronous as the Mach number increases beyond 0.90. This contradicts the trends predicted by linear flutter analyses and suggests that a change in the oscillation bounding mechanism occurs over the flight region examined.

Numerical Simulations of Membrane Wing Aerodynamics for Micro Air Vehicle Applications

Abstract: To gain insight into the aerodynamics of flexible wing-based micro air vehicles (MAVs), we study the threedimensional interaction between a membrane wing and its surrounding fluid flow. A nonlinear membrane structural solver and a Navier‐Stokes flow solver are coupled through the moving boundary technique and time synchronization. Under the chord Reynolds number of 9 × × 10 4 , the membrane exhibits self-initiated vibrations in accordance with its material properties and the surrounding fluid flow. The vortical flow structure, its effect on the aerodynamic parameters, and the implications of the membrane deformation on the effective angle of attack and flow structure are discussed. Nomenclature C D = drag coefficient CL = lift coefficient c = chord length c p = pressure coefficient D = drag Fpx = form drag Fpy = lift caused by pressure force Fτ x = drag caused by friction L = lift U = freestream speed u = chordwise velocity v =v ertical velocity x = chordwise distance from the leading edge Z = half-wing span z = spanwise distance from the root α = angle of attack

Propeller Performance at Low Advance Ratio

Abstract: At low advance ratios that are much smaller than the advance ratio where the maximum efficiency of the propeller is obtained, large portions of the blades' cross sections operate at stall conditions. Using two-dimensional nonrotating airfoil data to calculate the propeller's aerodynamic performance at low advance ratios results in large differences between the calculated and measured performance. It turns out that because of rotation the stall characteristics of the airfoil are changed because Coriolis effects delay the boundary-layer separation. Based on previous investigations associated with wind turbines that have been reported in the literature, a simple correction model is presented. It is a straightforward matter to implement this model in existing strip models. The agreement between the calculated and measured results, at low advance ratios, is significantly improved after the introduction of the new model. ROPELLERS are usually designed to operate optimally at a cer- tain design point, for example, the aircraft's cruise conditions. It is clear that propellers also operate at off-design flight conditions that can include takeoff, steep climb, etc. At off-design flight condi- tions, the efficiency of the propeller drops significantly. To avoid an operation at low efficiencies, propellers are often equipped with vari- able pitch mechanisms. Yet, there are many aircraft that do not have av ariable pitch mechanism because of price, weight, or reliability considerations. These included unmanned aerial vehicles (UAVs), ultralight, and low-priced general aviation aircraft. A fixed-pitch propeller that operates at flight speeds that are much lower than the cruise speeds, namely low advance ratios, suffers from a sharp drop in its efficiency. At low advance ratios, large portions of the blades are operating at stall conditions. Most of the literature that deals with propellers' performance does not include results for low advance ratios. Evans and Liner 1 present experimental results for propellers at low advance ratios, but they do not present calculated results for these cases. Similar results are also presented in Yaggy and Rogallo. 2 An attempt to use classical methods to calculate the performance of propellers at low advance ratios (advance ratios that are much lower than the advance ratio where maximum efficiency is obtained) exhibits large differ- ences between the calculated and measured thrust, while the same methods exhibit excellent agreement at higher advance ratios. The differences between the calculations and measurements increase as larger portions of the blades experience high cross-sectional an- gles of attack, beyond the stall limit. Thus, it becomes clear that there are problems in modeling the stall characteristics of cross sec- tions of propellers' blades. It turns out that the measured propeller's thrust is higher (sometimes much higher) than the thrust predicted by calculations where the stall characteristics of a two-dimensional nonrotating airfoil are used. Himmelskamp 3 was probably the first to investigate the influence of rotation on the stall characteristics of a rotating airfoil. He ob- served lift coefficients as high as three near the hub of a rotating fan blade. Although it seems to the authors that Himmelskamp's results were not applied in previous analyses of propellers, they were widely used, especially during the last 15 years, in the aerodynamic analysis

Assessment of comprehensive analysis calculation of airloads on helicopter rotors

Abstract: Blade section normal force and pitching moment were investigated for six rotors operating at transition and high speeds: H-34 in flight and wind tunnel, SA 330 (research Puma), SA 349/2, UH-60A full-scale, and BO-105 model (Higher-Harmonic Acoustics Rotor Test I). The measured data from flight and wind-tunnel tests were compared with calculations obtained using the comprehensive analysis CAMRAD II. The calculations were made using two free-wake models: rolled up and multiple trailer with consolidation models. At transition speed, there is fair to good agreement for the blade section normal force between the test data and analysis for the H-34, research Puma, and SA 349/2 with the rolled-up wake. The calculated airloads differ significantly from the measurements for the UH-60A and BO-105. Better correlation is obtained for the UH-60A and BO-105 by using the multiple trailer with consolidation wake model. In the high-speed condition, the analysis shows generally good agreement with the research Puma flight data in both magnitude and phase. However, poor agreement is obtained for the other rotors examined. The analysis shows that the aerodynamic tip design (chord length and quarter-chord location) of the research Puma has an important influence on the phase correlation.

Aeroelastic Stability Enhancement and Vibration Suppression in a Composite Helicopter Rotor

Abstract: An optimization procedure to 1) reduce the 4/revolution oscillatory hub loads and 2) increase the lag mode damping of a four-bladed soft-in-plane hingeless helicopter rotor is developed using a two-level approach. At the upper level, response surface approximations to the objective function and constraints are used to find the optimal blade mass and stiffness properties for vibration minimization and stability enhancement. An aeroelastic analysis based on finite elements in space and time is used. The numerical sampling needed to obtain the response surfaces is done using the central composite design of the theory of design of experiments. The approximate optimization problem expressed in terms of quadratic response surfaces is solved using a gradient-based method. Optimization results for the vibration problem in forward flight with unsteady aerodynamic modeling show a vibration reduction of about 15%. The dominant loads are the vertical hub shear and the rolling and pitching moments, which are reduced by 22-26%. The results of stability enhancement problem show an increase of 6-125% in the lag mode damping. At the lower level, a composite box beam is designed to match the upper-level beam blade stiffness and mass using a genetic algorithm which permits the use of discrete ply angle design variables such as 0, +or-45, and 90 deg, which are easier to manufacture. Three different composite materials are used for designing the composite box beam, thus, showing the robustness of the genetic algorithm approach. Boron/epoxy composite gives the most compact box beam, whereas graphite/epoxy gives the lightest box beam

Effect of camber on the aerodynamics of adaptive-wing micro air vehicles

Abstract: Four microair vehicle wind-tunnel models were built with 3, 6, 9, and 12% camber, all based upon the S5010-TOP24C-REF thin, cambered-plate airfoil. These models were tested in the Low Speed Wind Tunnel at angles of attack ranging from 0 to 35 deg and velocities of 5, 7.5, and 10 m/s, corresponding to mean aerodynamic chord Reynolds numbers of 5 × 10 4 , 7.5 × 10 4 , and 1 x 10 5 , respectively. Aerodynamic coefficients C L , C D , C M and lift-to-drag ratio (LID) were obtained and plotted vs angle of attack for all of the cambers at each velocity

Airbus, ONERA, and DLR Results from the Second AIAA Drag Prediction Workshop

Abstract: The results from DLR, Airbus, and ONERA from the Second AIAA Computational Fluid Dynamics Drag Prediction Workshop are presented. The lift, drag, and pitching moments are calculated for the DLR-F6 configuration at transonic flow conditions by solving the Reynolds-averaged Navier-Stokes equations on structured as well as on unstructured, hybrid grids.

Direct Aeroelastic Bifurcation Analysis of a Symmetric Wing Based on Euler Equations

Abstract: The application of a sparse matrix solver for the direct calculation of Hopf bifurcation points for the flexible AGARD 445.6 wing in a transonic flow modeled using computational fluid dynamics is considered. The iteration scheme for solving the Hopf equations is based on a modified Newton’s method. Direct solution of the linear system for the updates has previously been restrictive for application of the method, and the sparse solver overcomes this limitation. Previous work has demonstrated the scheme for aerofoil calculations. The current paper gives the first three-dimensional results with the method, showing that an entire flutter boundary for the AGARD 445.6 wing can be traced out in a time comparable to that required for a small number of time-marching calculations, yielding two orders of magnitude improvement when compared to the time-marching approach.