Showing papers in "Journal of Aircraft in 2006"

Abstract: Nomenclature Ar = rotor disk area CD = sectional drag coefficient CD0 = zero-lift drag coefficient Clα = lift-curve slope CP = power coefficient CPi = induced power coefficient CP0 = profile power coefficient CT = thrust coefficient c = chord length D = drag force D.L . = disk loading L = lift force m = mass P.L . = power loading SF = separated flow T = rotor thrust V = local wind velocity perceived by flap W = weight W f = final weight Wo = gross takeoff weight α = blade section angle of attack η = efficiency μ = dynamic viscosity ρ = air density σ = rotor solidity = flapping amplitude (peak to peak)

Abstract: The paper presents a theory for flight-dynamic analysis of highly flexible flying-wing configurations. The analysis takes into account large aircraft motion coupled with geometrically nonlinear structural deformation subject only to a restriction to small strain. A large motion aerodynamic loads model is integrated into the analysis. The analysis can be used for complete aircraft analysis including trim, stability analysis linearized about the trimmed-state, and nonlinear simulation. Results are generated for a typical high-aspect-ratio "flying-wing" configuration. The results indicate that the aircraft undergoes large deformation during trim. The flight-dynamic characteristics of the deformed aircraft are completely different as compared with a rigid aircraft. When the example aircraft is loaded sufficiently, the pair of complex-conjugate short-period roots merges to become two real roots, and the phugoid mode goes unstable. Furthermore, nonlinear flight simulation of the aircraft indicates that the phugoid instability leads to catastrophic consequences.

Abstract: A computational fluid dynamics (CFD) code and rotorcraft computational structural dynamics (CSD) code are coupled to calculate helicopter rotor airloads across a range of flight conditions. An iterative loose (weak) coupling methodology is used to couple the CFD and CSD codes on a per revolution, periodic basis. The CFD code uses a high fidelity, Navier‐Stokes, overset grid methodology with first principles-based wake capturing. Modifications are made to the CFD code for the aeroelastic analysis. For a UH-60A Blackhawk helicopter, three challenging level flight conditions are computed: 1) high speed, μ = 0.37, with advancing blade negative lift, 2) low speed, μ = 0.15, with blade‐vortex interaction, and 3) high thrust with dynamic stall, μ = 0.24. Results are compared with UH-60A Airloads Program flight test data. For all cases the loose coupling methodology is shown to be stable, convergent, and robust with full coupling of normal force, pitching moment, and chord force. In comparison with flight test data, normal force and pitching moment phase and magnitude are in good agreement. The shapes of the airloads curves are well captured. Overall, the results are a noteworthy improvement over lifting line aerodynamics used in rotorcraft comprehensive codes.

Abstract: This paper presents a novel optimization approach to the design of mechanisms in morphing aircraft structures. The layout of the mechanism and the location and number of actuators and pivots are determined by an extended formulation of a material-based topology optimization. The design problem is modeled within a coupled fluid-structure analysis framework to directly assess aerodynamic performance criteria while optimizing the overall mechanized system. The proposed methodology is illustrated through the design optimization of a quasi-three-dimensional section of an adaptive wing, where the approach is compared to a conventional two-step approach of first optimizing the aerodynamic shape for one or multiple flight conditions, and then finding the mechanism that leads to this shape. The comparison shows that the interactions between flow, structural deformation, mechanism, and actuator must be considered to find the optimal solution. The optimization approach presented allows direct consideration of these interactions at the expense of an increased computational burden.

Abstract: Further developments, applications, and assessments of the probabilistic two-phase aircraft wake-vortex model P2P are described. The wake-vortex model is applied to data of two field measurement campaigns accomplished at Tarbes airport, France. Measurements corroborate unambiguously the two-phase circulation decay anticipated by theory and predicted by P2P. Vortex age and descent speed are adjusted to match effects of axial wind and glide slope angle. Envelopes of vortex trajectories are expanded to consider tilting, stalling and rebounding wake vortices caused by axial- and crosswind shear. For probabilistic model output a choice between arbitrary degrees of probability is established and a stochastic prediction mode is introduced. In a deterministic scoring procedure, model perfomance is compared to the skill of another model. Probabilistic model performance is evaluated by the compilation of probability density distributions which relate wake vortex measurement data to the predicted envelopes.

Abstract: Two coflowjet airfoils with injection slot size differing by a factor of 2 are tested in a wind tunnel to study the effect of injection slot size. At the same angle of attack, the larger injection slot size airfoil passes about twice the jet mass flow rate of the smaller injection slot size airfoil. The smaller injection slot size airfoil is more effective in increasing the stall margin and maximum lift, whereas the larger slot coflow jet airfoil is more effective in reducing drag. To achieve the same lift, the smaller injection slot size airfoil has much less energy expenditure than the larger injection slot airfoil. A coefficient of jet kinetic energy is introduced, which appears to correlate well with the maximum lift and stall margin when coflow jet airfoil geometry varies. Both the jet kinetic energy coefficient and the momentum coefficient correlate well with drag reduction. No optimization of the airfoil configuration is pursued in this research, and the results indicate that there is a great potential for coflow jet airfoil performance improvement.

Abstract: Parallel hybrid-electric propulsion systems would be beneficial for small unmanned aerial vehicles used for military, homeland security, and disaster-monitoring missions involving intelligence, surveillance, or reconnaissance (ISR). The benefits include increased time on station and range as compared to electric-powered unmanned aerial vehicles and reduced acoustic and thermal signatures not available with gasoline-powered unmanned aerial vehicles. A conceptual design of a small unmanned aerial vehicle with a parallel hybrid-electric propulsion system, the application of a rule-based controller to the hybrid-electric system, and simulation results are provided. The two-point conceptual design includes an internal combustion engine sized for cruise speed and an electric motor and lithium-ion battery pack sized for endurance speed. A rule-based controller based on ideal operating line concepts is applied to the control of the parallel hybrid-electric propulsion system. The energy use for the 13.6 kg (30 Ib) hybrid-electric unmanned aerial vehicle with the rule-based controller during one-hour and three-hour ISR missions is 54% and 22% less, respectively, than for a four-stroke gasoline-powered unmanned aerial vehicle.

Abstract: Autopilot systems are capable of reliably following flight plans under normal circumstances, but even the most advanced flight-management systems cannot provide robust response to most anomalous events including in-flight failures. This paper describes an emergency flight-management architecture that can be applied to piloted or autonomous aircraft, with focus on the design and implementation of an adaptive flight planner (AFP) that dynamically adjusts its model to compute feasible flight plans in response to events that degrade aircraft performance. A two-step landing-site selection/trajectory generation process defines safe emergency plans in real time for situations that require landing at an alternate airport. A constraint-based search algorithm selects and prioritizes feasible emergency landing sites, then the AFP synthesizes a segmented trajectory to the best site based on postfailure flight dynamics. The AFP architecture is general for any failure situation; however, operational success is guaranteed only with accurate postfailure performance characterization and a trajectory generation strategy that respects degraded flight envelope boundaries. A real-time segmented trajectory planning algorithm and case study results are presented for total loss of thrust failure scenarios. This emergency is surprisingly common and necessitates an immediate approach and landing without a go-around option.

Abstract: Thispaperpresentsanadjointmethodfortheoptimumshapedesignofunsteadythree-dimensionalviscous flows. The goal is to develop a set of discrete unsteady adjoint equations and the corresponding boundary condition for the nonlinear frequency-domain method. First, this paper presents the complete formulation of the time-dependent optimal design problem. Second, we present the nonlinear frequency-domain adjoint equations for threedimensional viscous transonic flows. Third, we present results that demonstrate the application of the theory to a three-dimensional wing.

Abstract: This paper describes the use of full-span deforming control surfaces (including controlled chordwise cambering of the wing) to provide effective subsonic induced drag control with precise control of the spanwise lift distribution. Examples of these proposed controllers are smart, adaptive actuators for advanced unmanned air vehicle concepts. Reshaping the wing spanwise lift distribution with aeroelastic tailoring concepts is shown to reduce induced drag at high airspeeds. Active controllers such as conventional ailerons or leading-edge devices also reduce induced drag if they have a tailored spanwise deflection pattern; the required deflections of these surfaces depend on wing deformation and can be so large that they are not practical. Combining wing aeroelastic stiffness tailoring with active control surface design to create a control-friendly structure reduces induced drag and requires only small controller inputs. An exact solution for the actuator deflections to generate an elliptical lift distribution for the aeroelastic wing, for minimum induced drag, is discussed. A formal optimization problem is posed for cases in which multiple surfaces are used to control drag. This controller optimization solution is complicated because aeroelastic phenomena, such as control reversal, limit the effectiveness of some actuators at high speed.

Abstract: A computational fluid dynamics (CFD) model is coupled with a computational structural dynamics (CSD) model to improve prediction of helicopter rotor vibratory loads in high-speed flight. The two key problems of articulated rotor aeromechanics in high-speed flight-advancing blade lift phase, and underprediction of pitch link load-are satisfactorily resolved for the UH-60A rotor. The physics of aerodynamics and structural dynamics is first isolated from the coupled aeroelastic problem. The structural and aerodynamic models are validated separately using the UH-60A Airloads Program data. The key improvement provided by CFD over a lifting-line aerodynamic model is explained. The fundamental mechanisms behind rotor vibration at high speed are identified as: 1) large elastic twist deformations and 2) inboard wake interaction. The large twist deformations are driven by transonic pitching moments at the outboard stations. CFD captures 3-dimensional unsteady pitching moments at the outboard stations accurately. CFD/CSD coupling improves elastic twist deformations via accurate pitching moments and captures the vibratory lift harmonics correctly. At the outboard stations (86.5% radius out), the vibratory lift is dominated by elastic twist. At the inboard stations (67.5% and 77.5% radius), a refined wake model is necessary in addition to accurate twist. The peak-to-peak pitch link load and lower harmonic waveform are accurately captured. Discrepancies for higher harmonic torsion loads remain unresolved even with measured airloads. The predicted flap-bending moments show a phase shift of about 10 deg over the entire rotor azimuth. This error stems from 1, 2, and 3/rev lift. The 1/rev lift is unaffected by CFD/CSD coupling. The 2 and 3/rev lift are significantly improved but do not fully resolve the 2 and 3/rev bending moment error.

Abstract: An experimental investigation on aircraft landing gear noise is presented. The study includes systematic testing and data analysis using a full-scale Boeing 737 landing gear. The database covers a range of mean flow Mach numbers typical of landing conditions for commercial aircraft and various landing gear configurations, ranging from a fully dressed, complete gear to cleaner configurations involving only some parts of the complete gear. This enables the examination of noise radiation from various groups of the gear assembly and the derivation of functional dependencies of the radiated noise on the flow Mach number at various far-field directivity angles and on various gear geometry parameters. It is shown that the noise spectrum can be decomposed into three frequency components, namely, the low-, mid-, and high-frequency components, respectively, representing contributions from the wheels, the main struts, and the small details such as hoses, wires, cutouts, and steps. It is found that these different frequency components have different dependencies on flow parameters and gear geometry. Based on the spectral decomposition in the three frequency domains, normalized spectra are derived for all three components and a model for the overall sound levels is developed as a function of flow and geometry parameters.

Abstract: This paper introduces a morphing aircraft concept whose purpose is to demonstrate a new bioinspired flight capability: perching. Perching is a maneuver that uses primarily aerodynamics, as opposed to thrust generation, to achieve a vertical or short landing. The flight vehicle that will accomplish this is described herein with particular emphasis on its addition levels of actuation beyond the traditional aircraft control surfaces. The dynamics of this aircraft are examined with respect to changing vehicle configuration and flight condition. The analysis methodologies include an analytical and empirical aerodynamic analysis, trim and stability analyses, and flight simulation.Forthisstudy,theaircraft’smotionsarelimitedtothelongitudinalplaneonly.Specifically,cruiseandthe perching maneuver are examined, and comparisons are drawn between maneuvers involving vehicle reconfiguration and those that do not.

Abstract: A more practical analytical solution for the effects of wing twist on the performance of a finite wing of arbitrary planform has recently been presented. This infinite series solution is based on Prandtl's classical lifting-line theory, and the Fourier coefficients are presented in a form that depends only on wing geometry. Except for the special case of an elliptic planform, this solution shows that, if properly chosen, wing twist can be used to reduce the induced drag for a wing producing finite lift. A relation for the optimum twist distribution on a wing of arbitrary planform was presented. If this optimum twist distribution is used, the new solution predicts that a wing of any planform can be designed for a given lift coefficient to produce induced drag at the same minimum level as an elliptic wing having the same aspect ratio and no twist. In the present paper, results predicted from this new lifting-line solution are compared with results predicted from computational-fluid-dynamics (CFD) solutions. In all cases, the CFD solutions showed that the drag reduction achieved with optimum twist was equal to or greater than that predicted by lifting-line theory.

Abstract: Reported are the results of wind-tunnel tests conducted to evaluate aerodynamics characteristics of aircraft in formation flight. A vortex-lattice numerical scheme was used to investigate the effect of spatial offset (horizontal and vertical) between the leading and trailing wings. The wind-tunnel test configurations consisted of echelon, chevron, and in-line formations. Analysis of the data revealed that the spatial offset and the angle of attack of the leading wing had significant impact on the trailing aircraft. For some test conditions an increase in lift-to-drag ratio of the trailing aircraft was measured. Variation in C Lmax and/or α stall was observed as well. At higher angles of attack of the leading wing, the C L-α curve of the trailing aircraft was significantly altered.

Abstract: A novel scheme is presented for an iterative decambering approach to predict the post-stall characteristics of wings using known section data as inputs. The new scheme differs from earlier ones in the details of how the residual is computed. With this scheme, multiple solutions at high angles of attack are brought to light right during the computation of the residual for the Newton iteration. As with earlier schemes, multiple solutions are obtained for wings at high angles of attack and the resulting converged solution depends on the initial conditions used for the iteration. In general, the new scheme is found to be more robust at achieving convergence. Results are presented for a rectangular wing with two different airfoil lift curves and for a wing-tail configuration.

Abstract: The unsteady aerodynamics of a biomimetic inspired flapping-wing mechanism has been analyzed by performing detailed phase-locked diagnostics of its flow field. Flow visualization and particle image velocimetry results have shown the presence of a shed dynamic stall vortex that spans across most of the wing span. The shedding of this type of leading-edge vortex was accompanied by the formation of another leading-edge vortex before the first vortex reached the midchord, resulting in multiple shedding leading-edge vortices on the top surface of the wing during each wing stroke. A strong starting vortex was also formed at the trailing edge of the wing during the early part of its translational stroke. This vortex continuously gained strength from shed vorticity as the wing accelerated into its stroke. The starting vortex remained close to the trailing edge until the wing reached midstroke. A pair of vortices that continuously trailed from the root and tip of the wing were identified, both of which induced a significant downwash velocity over the wing surface. These trailed vortices were found to exhibit a contracting wake structure as they convected into the wake below the wing, consistent with an increase in slipstream velocity. The evolution of the tip and root vortex pair showed rapid diffusive characteristics with an increase in time (wake age).

Abstract: This paper analyzes the main contributors of infrared (IR) signature in a typical aircraft on a low-altitude mission. Various computational models are used to predict IR radiation from the aircraft. The bands within IR spectrum in which aircraft are susceptible to a typical IR-guided surface-to-air and air-to-air missile, for typical cases of tactical relevance, are identified. Lock-on range for aircraft against a typical missile is also computed. The feasibility of a low-altitude mission against a ground-based IR-guided threat is analyzed. The technique of emissivity optimization of aircraft rear fuselage skin, for reducing its infrared signature, is introduced and compared with other IR signature suppression techniques. The effectiveness of this technique in enlarging the safe flight envelope of aircraft, with respect to threat from heat-seeking missiles, for both surface-to-air and air-to-air missiles, is demonstrated. It is found that earthshine reflected off the aircraft surface plays a crucial role in the effectiveness of this technique against a surface-to-air missile (SAM) in 8‐12 μm band. Nomenclature A = area, m 2 H = spectral irradiance, μW/μm · m 2 h = aircraft altitude, km I = spectral radiant intensity, W/Sr · μm · m 2 J = spectral radiance: comprising emission and earthshine, W/Sr · μm · m 2 L = length, m M = Mach number N = number of discretized elements NEI = noise equivalent irradiance, W/m 2 Rma = distance separating missile and aircraft, km

Abstract: This article summarizes the results of an optimization study of a microvortex generator flow control in an inlet. Five parameters optimization was carried out using the classical design of experiment method. Two main objectives were in focus: first, to develop the methodology and skills necessary to conduct a design of experiment optimization study in area of the flow control; and second, to develop the procedures which would be used during design of vortex generator flow control in inlets. New information about the dependency of the vortex generator flow control in inlet on its geometrical parameters were obtained. Several interesting configurations were located. The parameters of optimal settings were then used to set up the vortex generator installation in a generic inlet.

Abstract: This research was supported by the Korea Airspace Research Institute (KARI) for a program of development of the CNS/ATM system for the next generation. We truly appreciate their financial support.

Abstract: Flow around an airfoil with a deploying microtab device has been numerically simulated by solving the unsteady turbulent compressible Navier-Stokes equations with the OVERFLOW-2 solver. Using a Chimera/overset grid topology, microtabs were placed at 95% ofchord of a symmetric NACA 0012 airfoil. Microtab heights on the order of 1% of chord, deployed on the order of one characteristic time unit were utilized. The unsteady effects of tab deployment time, deployment height, and freestream angle of attack on aerodynamic responses were also investigated. Validation studies with experimental results for static deployed microtabs and a dynamically deployed spoiler were also performed to ensure accurate temporal and spatial resolution of the numerical simulations.

Abstract: Miniflaps at the trailing edge of airfoils, that is, Gurney flaps, change the Kutta condition and thereby produce higher lift. Unfortunately, because of the flow separation downstream of such trailing edges, the drag also increases. Investigations are described with the aim to stabilize the wake flow to achieve drag reduction. When hot-wire anemometry is used, a tonal component in the spectrum of the velocity fluctuations downstream of the Gurney flap is shown. This points to the existence of a von Karman vortex street. Modifications of the Gurney flap can reduce this flow instability, which results in a drag reduction. Trailing-edge modifications, such as slits or holes in Gurney flaps and vortex generators, were tested in experiments. The experiments were carried out using straight wings and a swept wing at a Re = 1 × 10 6 At lower angles of attack of the airfoils with geometrical modifications a drag reduction was observed. This drag reduction was determined through force measurements. The flowfield behind the Gurney flaps was also investigated numerically, using methods based on Reynolds averaged Navier-Stokes and detached eddy simulation

Abstract: An experimental delta-wing/store model with freeplay has been designed and tested in the Duke wind tunnel. The wing structure is modeled theoretically using von Karman plate theory that accounts for geometric strain- displacement nonlinearities in the plate wing structure. A component modal analysis is used to derive the full structural equations of motion for the wing/store system. A linear three-dimensional time-domain vortex lattice aerodynamic model including a reduced-order model aerodynamic technique and a slender-body aerodynamic theory for the store are also used to investigate the nonlinear aeroelastic system. The effects of the freeplay gap, the span location of the store, and the initial conditions on the limit-cycle oscillations (LCO) are discussed. The correlations between the theory and experiment are good for the smaller LCO amplitudes, that is, for flow velocities slightly higher than the flutter velocity, but are not good for the larger LCO amplitudes, that is, higher flow velocities. The theoretical model needs to be improved to determine LCO response for larger-amplitude motions.

Abstract: The emerging flight-by-wire and flight-by-light technologies increase the possibility of enabling and improving aircraft design with excellent handling qualities and performance across the flight envelope. As a result, it is desired to take into account the dynamic characteristics and automatic control capabilities at the early conceptual stage. In this paper, an integrated control-configured aircraft design sizing framework is presented. It makes use of multidisciplinary design optimization to overcome the challenges which the flight dynamics and control integration present when included with the traditional disciplines in an aircraft sizing process. A commercial aircraft design example demonstrates the capability of the proposed methodology. The approach brings higher freedom in design, leading to aircraft that exploit the benefits of control configuration. It also helps to reduce time and cost in the engineering development cycle.

Abstract: A unified aeroelastic and flight dynamic formulation is sought to take into account the influence of aeroelastic effects on the flight dynamic behavior of the whole aircraft in a format fully compatible with the aeroelastic, flight dynamics, and automatic control disciplines. By allowing the inclusion of gravity-related terms, vertical acceleration-related aerodynamic stability derivatives, and lift and drag forces due to forward-velocity perturbations into the rational function approximation matrices, the traditional quasi-steady flight dynamic equations of motion are fully recovered. Closed-form solutions are presented for translational and rotational degrees of freedom in the aeroelastic model. The General Atomics-Aeronautical Systems (GA-ASI) Predator® unmanned aerial vehicle is used to numerically demonstrate the unified aeroelastic modeling framework. The results indicate that this approach reproduces, with a high degree of fidelity, the underlying quasi-steady flight dynamic model when no elastic modes are included in the aeroelastic model. Comments are provided to determine the approximate number of elastic modes that need to be included in the aeroelastic model to accurately model its flight dynamic behavior.

Abstract: The presented work addresses the need to integrate cost into the early product definition process as an engineering parameter. The methodology developed is generic and fundamental in developing causal predictions of manufacturing cost that are driven by the design parameters which give rise to the main elements of that cost relative to process capability. The manufacturing cost modelling is original and relies on a genetic-causal method of 1) classifying the generic cost elements that are linked to particular genetic identifiers relating to materials, processes and form; 2) developing causal parametric relations that link those genetic identifiers to a parent manufacturing cost. The application studied is a fuselage panel that is typical to commercial transport regional jets. Consequently, a semi-empirical numerical analysis using ESDU reference data was coupled to model the structural integrity of thin-walled structures with regard to material failure and buckling: skin, stringer, flexural and inter-rivet. The optimisation process focuses on Direct Operating Cost (DOC) as a function of acquisition cost and fuel burn. It was found that the ratio of acquisition cost to fuel burn was typically 4:3 and that there was a 10% improvement on the DOC for the minimal DOC condition over the minimal weight condition; due to the manufacturing cost saving from having a reduced number of larger- area stringers and a slightly thicker skin than preferred by the minimal weight condition. It is also noteworthy that the minimal manufacturing cost condition was slightly better than the minimal weight condition, which highlights the key finding: the traditional minimal weight condition is a dated and sub-optimal approach to airframe structural design.

Abstract: Flowfield measurements were carried out on the upper surface of a GLC-305 airfoil configured with glaze and rime ice-shape simulations. The mean and root-mean-square fluctuation of the streamwise velocity were measured using a split-hot-film probe at several chordwise locations. These data were taken at three different angles of attack preceding stall for each iced-airfoil configuration at Reynolds numbers of 3.5 × 10 6 and 6.0 x 10 6 with Mach numbers of 0.12 and 0.21. The velocity measurements confirmed the presence of a large separation bubble downstream of the ice shapes. The separation bubbles for the glaze ice configuration were much larger than those for the rime ice case, resulting from the differences in the ice horn geometry. Other than the differences in size, the integral boundary-layer characteristics were very similar. Changes in Reynolds number did not significantly affect the separation bubble characteristics. However, a larger Mach number did result in a slightly larger separation bubble for the glaze ice case at α = 6 deg. The root-mean-square velocity distributions had peak values in the separated shear layer, downstream of transition, that compared well with previous work.

Abstract: The electroimpulse de-icing system of an aircraft wing leading edge is investigated through the development of a methodology for the numerical simulation of the electroimpulse de-icing process. The principle of electroimpulse de-icing is that the ice is removed due to the leading edge local mechanical vibration, which is induced by an electromagnetic pulse. The numerical simulation methodology is based on a nonlinear transient three-dimensional stress analysis of the ice-covered wing, combined to a de-icing criterion that takes into consideration the tensile and shear stresses at the ice-skin interface. The developed methodology is verified on de-icing experimental tests of an aluminum plate. Afterwards, the methodology is applied to the prediction of de-icing of an aircraft wing leading edge. The dominant process parameters are determined to be the coil number and position, the ice thickness and coverage, the radius of wing curvature, and the electroimpulsive load amplitude. A parametric study is performed to determine the influence of the process parameters on the system effectiveness, defined as the percentage of the de-iced surface over the total leading edge surface. From the results of the parametric study, the possibilities of reducing the weight and energy consumption of the electroimpulse de-icing system can arise.

Abstract: The formation and near-field development of a wing-tip vortex under the influence of freestream turbulence were examined using flow visualization and hot-wire anemometry. A low turbulence freestream as well as two cases of grid turbulence with different intensities and length scales were considered. In all cases, the tip vortex was found to form from three smaller vortices, but the turbulence in its core was found to intensify with increasing freestream turbulence. The vortex trajectory was found to be unaffected by freestream turbulence, but the wing wake that was rolling up around the vortex was observed to have a curvature that decreased as freestream turbulence increased. The mean axial velocity distribution in the low-turbulence case was neither jetlike nor wakelike but had an annular shape. Time-averaged velocity profiles measured in the turbulent freestream cases were wakelike, and it was inferred that the instantaneous profiles would be significantly affected by vortex meandering. Mean circumferential velocity distributions in the vortex core displayed self-similar developments in all cases examined. Finally, it was found that the apparent diffusion in the shear layer shed from the wing increased with increasing freestream turbulence.

Abstract: The results of a Rolls-Royce sponsored programme of design to cost research are outlined. The role of cost modelling within the design process for the development of a civil gas turbine engine is outlined. The novel application of a generic financial modelling tool to an engineering cost estimating problem is demonstrated. This use of this tool to capture and disseminate costing knowledge is described and the use of modular library elements to develop cost models is shown. A prototype systems for the creation of an elegant cost model structure to enable direct integration with a CAD representation of a part and the integration of the costing capability within an automated design search and optimization environment is described