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The design and performance quantification of four Vertical Take-Off and Landing (VTOL) architectures for a canard-type aircraft configuration are presented. The aero-structural sizing of the canard configuration and the sizing procedure for the proposed VTOL configurations are described and discussed. The proposed VTOL architectures are based on a range of rotor distances to the centre of gravity, quad- and tri-rotor configurations, retractable front rotors and tilt rear rotors. The aerodynamic performance, total installed power and VTOL system mass were modelled and experimentally validated. The results show that a fully exposed VTOL system penalises the Lift-over-Drag (L/D) ratio significantly relative to a clean configuration. The VTOL system mass can be reduced by up to 32% by using a tilt tri-rotor configuration when compared with an equidistant quad-rotor+pusher configuration. The fraction of installed power usable for forward flight can be increased by up to 80% with a tilt configuration. For the proposed mission, the range can be significantly increased if a tri-rotor tilt configuration is adopted in place of an equidistant quad-rotor+pusher configuration.

The first-approach techniques for analytical evaluation of the airfoil shape and calculation of the aerodynamic coefficients for a box-wing aircraft are suggested. The reliability of the calculation results was estimated using FLZ code (the vortex lattice method). The comparison showed satisfactory convergence of coefficients for the clean configuration and some discrepancy for the take-off and landing configuration.

For flight safety reason, a better understanding of the effect of ice accretion on the aerodynamic performance is required. The ice shape used in this study is holding ice with one or two ice horns, which accretes on the unprotected surface of the aircraft. By using the verified numerical simulation method, it can be found that due to the effects of holding ice, the stall angle of the aircraft is significantly reduced and lift has a severe penalty. Besides, the pitching moment characteristics are deteriorated, which is a threat to flight safety. The main reason for the pitch up of aircraft with holding ice is the decreases in pitch-down moment of the horizontal tail affected by the iced wing and the increases in pitch-up moment provided by the wing. To avoid this, a method is proposed to improve the aerodynamic performance of iced high-lift configuration, which is shortened inner slat. It can be seen that the obvious improvement in pitching moment characteristics is at the expense of the slight decrease in lift for the iced high-lift configuration. Besides, this method has little effect on aerodynamic performance of the clean configuration.

The Distributed Electric Propulsion(DEP) UAV has the advantages of high efficiency, low noise, low emission and high control redundancy. Especially, it is more difficult about dynamic modeling and attitude controller design due to its Propulsion Airframe Integration(PAI) for DEP UAV with vectorable power section and the V/STOL capability. In this paper, aerodynamic data of different state points in the clean configuration of DEP UAV is caculated, and the nonlinear aerodynamic model is established by cubic spline interpolation method. Then, the wind tunnel data of the power section are analyzed, and the model for PAI is established. Finally, Linear Active Disturbance Rejection Control(LADRC) controller combined fal-function is designed for attitude control of nonlinear model of clean configuration aircraft, and PAI is added as controller disturbance in full speed envelope flight simulation. Taking pitch angle control as an example, compared with PID algorithm and LADRC, the modified LADRC method in this paper is not only more robust to PAI effects, but also better to the dynamic response of the system, which lays a foundation for S/VTOL of DEP UAV.

Results are presented on the freejet testing of a 75%-scale replica of the HIFiRE 7 scramjet. The HIFiRE 7 scramjet flowpath includes a two-dimensional forebody, a rectangular-to-elliptical shape t...

The purpose of this investigation is determinate the effects of propeller on wing aerodynamic, both for a propeller mounted in the middle of the wing, and for tip mounted propeller. Especially, it is investigated how a tip-mounted propeller can decrease wing induced drag, and how distributed propulsion can increase the high-lift aerodynamic. Analyses are carried out using a Virtual Disk Model on CFD software, showing a good agreement comparing numerical results with experimental data obtained by previous works. Wing tip engine with propeller, has been employed on a general aviation aircraft wing with an installed thrust to accomplish with cruise performance, reducing the induced drag. Distributed propeller engines on the wing allows improving of low speed performance, increasing the aircraft lift coefficient. Induced drag can be reduced of about 2-3% a low cruise lift coefficient, until 8-10% at relative high cruise lift coefficient. Maximum achievable lift coefficient could be increased of about 20-30% in clean configuration, and more than 50% in flapped configuration.

The development and comparison of aerodynamic simulation models for the inclusion in an open-loop stability & control model based on both experimental and numerical data are presented. Specifically, the work is focused on the effectiveness of the Kestrel simulation tools in replicating the aerodynamic coefficient data from wind tunnel experiments of a generic lambda wing configuration with four trailing edge control surfaces. Simulation conditions correspond to the angles of attack in range of − 5 ∘ to 10∘ at assumed cruise Mach number of 0.7. These simulations include a clean configuration with no control surface deflection and when inboard/outboard control surfaces are deflected − 10 ∘ , − 5 ∘ , 5∘, and 10∘. Single and overset grids are used to simulate the control surface deflections. Predicted aerodynamic data are collected in form of look-up tables and are run in a Matlab/Simulink program to determine stability and control properties of the vehicle. The vehicle stability and control characteristics are then compared with the results of running experimental coefficient values through the same program. The results show that numerical predictions of the test case considered depend heavily on the geometry details and numerical setup, e.g. turbulence modeling, grid refinement, integration scheme, time step, etc. and therefore do not match experimental data at all tested conditions. In addition, the stability and control investigation of the vehicle using these data confirms that the differences between simulations and measurements could lead to significant changes in the evaluation of vehicle's stability and control characteristics. The simulation predictions of this work, therefore, cannot fully replace the need for experiments to determine the vehicle's stability and control, however they help to provide insight into flow physics and to direct/make best use of wind tunnel experiments.

Purpose Realistic data bank of aerodynamic and stability derivatives is still missing for hybrid buoyant aerial vehicles. Such vehicles take-off and land similar to an aircraft with their partial weight balanced by the aerostatic lift. The purpose of this paper is to use wind tunnel testing for a better understanding of the aerodynamic and static stability behavior of such vehicles. Design/methodology/approach The effect of wing on the aerodynamic and static stability characteristics of a clean configuration hybrid buoyant is analyzed. The free stream velocity is 20 m/s, and ranges of angle of attack and side slip angle are from −8° to 12° and ±16°, respectively. Data are corrected to account for the effect of strut interference and zero load condition. The maximum blockage of the model with respect to the cross-section area of the test section is about 2.7 per cent. Findings A hybrid model manufactured by using wood and metal is an optimum solution with less number of parts. The vehicle is statically, longitudinally and directionally stable. Wings designed to fulfill the partial requirement of lift contribute significantly to counter the huge moment generated by the voluminous hull for centre of gravity location ahead of the leading edge of the wing. Research limitations/implications There are number of manufacturing constraints for scaling down a model of a hybrid buoyant aerial vehicle configuration. Specially, the thickness of the wing limits the testing envelop of angle of attack and free stream velocity. Practical implications The data presented here are a preliminary guide for further work on larger size models. The data may also be used to build and perform flight tests on small full-scale instrumented models and to obtain flight dynamics data. Originality/value The estimated aerodynamic and stability derivatives and slopes can be utilized in future for multidisciplinary design.

The aim of this work is the development of a methodology to predict lift characteristics for transport aircraft in the whole flight envelope, useful in the preliminary aircraft design stage. The purpose is an attempt to improve the classical methodologies for wing load distribution and lift prediction, applicable to both clean and flapped configuration. This has been obtained considering the airfoils’ aerodynamic characteristics until stall and post-stall conditions during the process, and modifying 2D characteristics in the case of high-lift devices to take into account 3D effects introduced by the devices themselves. The method is a modification of standard vortex-lattice procedures which are capable of predicting wing aerodynamic characteristics. As regards the clean configuration, the enhanced method works by integrating airfoil characteristics, whereas as far as the high-lift devices’ effect is concerned, the improved method works by substituting clean airfoil aerodynamic characteristics for the flapped aerodynamics ones, and introducing a correction to evaluate the 3D effects induced by the high-lift devices’ geometrical discontinuities. The methodology is explained separately for these two configurations. The results of the developed method have been compared with CFD and experimental data showing good agreement, making available a fast and reliable method useful in preliminary aircraft design phase.

Controlling the shape of the leading edge of the wing has great potential both in terms of aerodynamic efficiency (costs and pollution abatement) and high lift generation. Theoretically, alteration of the leading edges geometry (local and global curvature) should produce a smooth and effective configuration, assuring a hyperlift generation equivalent to traditional devices (e.g., slats) during take-off and landing, and preserving (through clean configuration) the laminar flow. Despite the aforementioned benefits, the morphing of the leading edge poses different technical issues, ranging from aerodynamics to mechanics and control. Because of the remarkable influence of its geometry on the aerodynamic field of the entire wing section, achieving an appropriate morphed shape in the leading edge is even more critical than for other zones of the wing. Due to laminar flow requirements, no asperities (curvature discontinuities or gaps) are allowed on the external surface. At the same time, the actuated shape must guarantee the necessary peak depression, minimizing any separation of the flow. This latter aspect implies a dedicated geometry characterized by a large global curvature, known as a “droop nose.” Mechanically, applying a droop nose geometry requires particular characteristics of some elements. For example, the skin has to be flexible enough to enable the droop nose effect, but must also withstand large aerodynamic loads, guaranteeing maintenance of the prescribed morphed shape. Parameter tolerances become even tighter in the context of airworthiness requirements, as applied to conventional leading edges (e.g., hail and bird impact). This section provides an overview of the state-of-the art morphing technologies suitable for leading edges, with specific reference to the most recent applications developed and matured in the SADE and SARISTU European projects. Then, an advanced, conceptual design example of droop nose architecture (conceived by CIRA within SADE) will be illustrated. Finally, the manufacture of a full-size demonstrator, and its lab validation, will be outlined.

International audience; The next generation of agile unmanned combat aerial vehicles has put focus on the stability and control of low observable platforms. Without the traditional fin for yaw control combined with delta wings, a challenging task has been presented to the research community. The NATO research group AVT-201 was formed to meet this challenge and build on previous knowledge gained on the complex vortical flow of rounded leading-edge delta wings. This paper presents two numerical studies made by the Swedish Defence Research Agency (FOI) and the French Aerospace Lab (ONERA) on traditional trailing-edge control and a more unconventional type of leading-edge flap control. Results from the FOI-developed flow solver Edge, using trailing-edge control, on a low observable 53 deg swept lambda wing model are compared to experimental data. The effect of leading-edge control is investigated by ONERA with a slightly smaller-scaled model, using the in-house code elsA, and validated by experimental data. Additional static and dynamic derivatives are evaluated by ONERA for the clean configuration.

CESAR focuses on small-size commercial aircraft providing manufacturers with an enhanced ability needed to become fully competitive in the world market. The objective is to build up a new development concept for this aircraft category and to improve selected technologies enabling a significant reduction of the time-to-market and lowering the overall development, operation and maintenance costs, while considering safety, passenger comfort and environmental impact. The CESAR aspires to provide technologies and knowledge for advanced wing, competitive and environmentally acceptable propulsion unit and new technologies for selected aircraft systems to reduce aircraft operating costs and improve safety. 1. WIND TUNNEL MODEL The AC2 model was built within the CESAR project according to the geometric and aerodynamic specifications for the CESAR Aircraft 2 turbulent wing and high lift devices. The AC2 model is designed to collect data for the operative envelope of the aircraft (for a clean configuration) in the INCAS trisonic wind tunnel. 2. MODEL DESCRIPTION The model scale was established at ks = 0.06. The AC2 model is designed for measurements of aerodynamic forces and moments by using the 2.00 MK XXVI TASK six components internal strain gauge balance. The AC2 model is built in a modular design that allows easy installation in INCAS trisonic wind tunnel on an existing U-shaped sting of a lenticular cross section. The sting is shown in figure 1 below

Flutter is a rapid self-feeding motion which is caused by the interaction of aerodynamic, structural and inertial forces. Flutter can cause major damage on aircraft structure which can lead to fatal accident in aviation. Several methods have been evolved to avoid the flutter phenomena occur during the flight envelope of aircraft design. On this study, method was developed by Indonesian Aerospace which consist of Finite Element Method (FEM) analysis, Ground Vibration Test (GVT), and Wind Tunnel Flutter Test (WTT). Based on the study, FEM have similar results toward to Wind Tunnel Flutter Test conjunction the clean configuration of N219 aircraft half wing model.

HyPlane is the Italian aerospaceplane proposal targeting, at the same time, both the space tourism and point-to-point intercontinental hypersonic flights. Unlike other aerospaceplane projects, relying on boosters or mother airplanes that bring the vehicle to high altitude, HyPlane will take off and land horizontally from common runways. According to the current project, HyPlane will fly sub-orbital trajectories under high-supersonic/low-hypersonic continuum flow regimes. It can go beyond the von Karman line at 100 km altitude for a short time, then starting the descending leg of the trajectory. Its aerodynamic behavior up to 70 km have already been studied and the results published in previous works. In the present paper some aspects of the aerodynamic behavior of HyPlane have been analyzed at 80, 90 and 100 km. Computer tests, calculating the aerodynamic parameters, have been carried out by a Direct Simulation Monte Carlo code. The effects of the Knudsen, Mach and Reynolds numbers have been evaluated in clean configuration. The effects of the aerodynamic surfaces on the rolling, pitching and yawing moments, and therefore on the capability to control attitude, have been analyzed at 100 km altitude. The aerodynamic behavior has been compared also with that of another aerospaceplane at 100 km both in clean and flapped configuration.

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.

This paper reviews the research on Gurney flaps and related high lift trailing edge devices. It investigates aerofoil performances at Reynolds numbers Re ≅ 105 and below, both with the clean configuration and various Gurney flap sizes. The device height is optimised, and a semi-empirical formula linking flap height to free stream speed and aerofoil chord is proposed. The analysis shows that the optimal size of the device is always below the boundary-layer thickness at the trailing edge. Discussion of results includes analysis of hysteresis loops occurring in the L/D performances. These are mostly due to large changes in drag and small changes in lift, which occur when the aerofoil is restored to the reference angle of attack.

An experimental wind-tunnel investigation was undertaken to determine the effects of Gurney flaps on a 40-deg cropped nonslender delta wing at a chord Reynolds number of 250,000. In the experiment, the height of the Gurney flaps was varied from 0.01C to 0.05C, and the sideslip angle of the model was selected as 0, 5, 10 and 20 deg. In addition, the 0.05C Gurney flap was serrated with different heights of 0.01C to 0.05C separately. In comparison with the baseline clean configuration results, it was found that the model with plate Gurney flaps can indeed increase the lift-to-drag ratio at moderate-to-high lift coefficients for the wing, and the greatest increment was obtained for the 0.01C Gurney flap. The effect of Gurney flap on the increment of lift-to-drag ratio tends to be not significant with the increase of sideslip angle. Moreover, the 0.05C serrated Gurney flap provides the best performance among the serrated Gurney flaps.

The paper presents the influence of ground distance on aerodynamic characteristics of the flow over NACA 4412 multi-element airfoil in clean configuration. The angle of deflection was kept constant throughout the computations. Finite volume method was used for the grid generation and the numerical calculation. Further investigation suggests the best range of angle of attack where the aerodynamic performance of the airfoil increases in the presence of ground. Furthermore, the condition of flow separation jeopardizing the airfoil performance appeared. It was found that at certain range of angle of attack, the cushion of high pressure at the pressure side of the airfoil is insufficient to compensate the influence of the separation. This work shed insight on the important parameters that need to be taken into account in the operational of an airplane.

This paper details a 2D numerical and experimental investigation on the change in flow dynamics when a Fowler flap gap and overlap are varied with respect to the trailing edge of the main wing at zero angle of attack. It was found that a commercially available CFD solver (Fluent) in RANS mode was in reasonable agreement with the experiment for the clean configuration only. For the landing configuration consisting of a Fowler flap deployed and deflected to 40◦, the CFD could not accurately compute the complex flow field. Much of the inaccuracy stemmed from the inability of CFD to accurately compute the flow in the gap. When the flow based on the experiment was imposed as a boundary condition, the computed flow field had a reasonable agreement with the experimental results.

A hinge moment based stall prediction method originally designed using 2D airfoil data is extended to a 3D wing. Experimental hinge moment measurements were obtained for a NACA 3415 semispan wing model in a clean configuration and with various simulated leading-edge contaminants. The effects of the simulated contamination configurations on the 3D wing and approximate 2D sectional performance are presented. These unsteady hinge moment data were processed through three detector functions to provide envelope protection for the clean and contaminated wing. The envelope protection advisories predicted by the detector functions are shown to be effective in predicting the angle-ofattack boundary of the normal flight envelope across various flap setting configurations and contamination configurations for the wing. Moreover, the use of the envelope protection system on data obtained during an unsteady pitch maneuver suggests that the envelope protection system may also function effectively in a dynamic environment. These results establish the viability of using this method on a flight vehicle.

In a joint effort ONERA and DLR have assessed the interference effects of a generic twin-engine transport aircraft configuration equipped with throughflow nacelles. Both experimental and theoretical investigations were performed. The results showed that the installation of the nacelles had substantial effects on total forces at cruise Mach number. With respect to the clean configuration, total lift decreased at constant incidence, and total drag increased at constant lift. A detailed analysis of the flowfield revealed that for transonic flows the installation of the nacelle influenced the complete upper wing surface. On the lower wing surface the flow was considerably accelerated due to the propulsion system, and a supersonic region terminated by a shock wave appeared. The nacelle pressure distribution was also substantially influenced by interference effects. Comparison of experimental and numerical results showed that the solution of the Euler equations is well capable of simulating the interference effects.

Computstional Fluid Dynamics is employed to predict aerodynamic coefficients for the second-stage configuration of a missile defense target missile. Models consist of the clean vehicle with fins and two additional models that include rail launch lugs and externally mounted rear facing camera pods. Analyses are performed to estimate wind tunnel test measurement requirements and to assist in planning of the wind tunnel test matrix. Postwind tunnel predictions for the clean configuration show excellent comparison to reduced test data. Both predictions and test measurements indicate that the protuberances contribute an additional axial drag to the clean missile configuration of over ten percent. I. Introduction HE Missile Defense Agency has developed a cost-effective target that can be used to assess missile defense capabilities. The Aegis Readiness Assessment Vehicle-Group C (ARAV-C) is a member of this target vehicle class. This paper documents the prediction of aerodynamic coefficients using CFD models for the ARAV-C second stage. Predictions are made for several points along a proposed flight trajectory as a pre-wind tunnel exercise to estimate effects required for measurement and to assist in planning of the test matrix. Post-wind tunnel predictions are made and compared to reduced test data. Favorable comparison resulted in additional computations requested to extend the wind tunnel data to higher Mach numbers for use in 6DOF simulations. Three separate models are generated: the clean vehicle with fins; a second with the addition of two externally mounted camera pods with no launch lugs; and a third that includes the addition of launch lugs with no camera pods. Predictions using the clean configuration and the configuration with camera pods at four roll angles and eight Mach numbers with angle of attack (AOA) set at 3° are made. Additional computations are made using all three configurations. Nineteen wind tunnel comparison computations are completed. A total of 131 3D viscous CFD computations are performed.

An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5 percent scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from alpha = -5deg to 85deg and beta = -45 deg to 45 deg at a Reynolds number of 0.24 x10(exp 6) and Mach number of 0.06. The 3.5 percent scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5 percent scale GTM. The addition of the large, glaze-horn type ice shapes did result in an increase in airplane drag coefficient but had little effect on the lift and pitching moment. The lateral-directional characteristics showed mixed results with a small effect of the ice shapes observed in some cases. The flow visualization images revealed the presence and evolution of a spanwise-running vortex on the wing that was the dominant feature of the flowfield for both clean and iced configurations. The lack of ice-induced performance and flowfield effects observed in this effort was likely due to Reynolds number effects for the clean configuration. Estimates of full-scale baseline performance were included in this analysis to illustrate the potential icing effects.

The NASA Johnson Space Center (JSC) in Houston, Texas is the home of the NASA WB-57 High Altitude Research Program. Three fully operational WB-57 aircraft are based near JSC at Ellington Field. The aircraft have been flying research missions since the early 1960's, and continue to be an asset to the scientific community with professional, reliable, customer-oriented service designed to meet all scientific objectives. The NASA WB-57 Program provides unique, high-altitude airborne platforms to US Government agencies, academic institutions, and commercial customers in order to support scientific research and advanced technology development and testing at locations around the world. Mission examples include atmospheric and earth science, ground mapping, cosmic dust collection, rocket launch support, and test bed operations for future airborne or spaceborne systems. During the return from a 6 hour flight, at 30,000 feet, in the clean configuration, traveling at 175 knots indicated airspeed, in un-accelerated flight with the auto pilot engaged, in calm air, the 2-man crew heard a mechanical bang and felt a slight shudder followed by a few seconds of high frequency vibration. The crew did not notice any other abnormalities leading up to, or for the remaining 1 hour of flight and made an uneventful landing. Upon taxi into the chocks, the recovery ground crew noticed the high frequency long wire antenna had become disconnected from the vertical stabilizer and was trailing over the left inboard wing, and that the left engine upper center removable cowling panel was missing, with noticeable damage to the left engine inboard cowling fixed structure. The missing cowling panel was never recovered. Each engine cowling panel is attached to the engine nacelle using six bushings made of 17-4 PH steel. The cylinder portions of four of the six bushings were found still attached to the aircraft (Fig 1). The other two bushings were lost with the panel. The other four bushings exhibited ratchet marks (multiple fatigue origins) which initiated in the sharp radius of the flange/cylinder fillet and were observed 300 degrees around the flange perimeter (Fig 2-3). Low stress, high cycle fatigue (HCF) was observed on the fracture surfaces of all four bushings (Fig 4). To improve the cowling panel joint design and enable return to flight, new cowling bushings with thicker flanges and a larger machined flange/cylinder fillet radius were installed on all cowling panels. In addition, a spacer was added to the joint to achieve the proper stack tolerance. Finally, a time change requirement for all cowling bushings was instituted.

This wind tunnel research, sponsored by the Natick Soldier Research, Development and Engineering Center (NSRDEC), investigates the flow behind a 1/84 scale model of the C-17 Globemaster III in a clean configuration with troop and cargo doors closed. A pressure rake swept through a region perpendicular to the aircraft’s line of flight generated two-dimensional maps of the aircraft wake at fixed distances away from the cargo door. Recirculation and separation in this region can adversely influence airdrop operations, and a detailed understanding of the flow-field is essential to developing engineering and operational approaches to mitigate their effects and increase mission success and safety. The tests covered several angles of attack and Mach numbers. The results demonstrated that the size and intensity of the wake near the cargo door were directly affected by the angle of attack of the aircraft, but the aircraft airspeed had a much smaller affect on the pressure recovery fraction data. At higher angles of attack, the aft section of the aircraft experienced more uniform free-stream flow that was less turbulent than at lower angles of attack. The lift and drag force data obtained, in conjunction with the pressure flow-field surveys will be used to validate future Computational Fluid Dynamics (CFD) studies. A statistical analysis of the velocity field near the body of the aircraft suggests the presence of vortices; however, the results are inclusive and will be investigated further with the aid of CFD analyses. The uncertainty in the wake measurements was less than 2%.

In order to meet the anticipated future demand for air travel, DLR's Institute of Flight Guidance is investigating Distributed Air-Ground Traffic Management (DAG-TM) for arrivals in the Extended TMA. Based on the exchange of information between aircraft and ATC via data link, a trajectory based traffic management can take into account user preferred trajectories as well as make use of highly accurate prediction of aircraft movements. This contribution will focus on the airborne part - in particular on the 4D FMS functions for planning and flying various types of CDA approaches. The planning functions to compute approach profiles ranging from 2 degree (clean configuration) up to steep slopes (more than 5 degree) with early configuration and glide path intersection from above will be presented. Furthermore, functions to compensate deviations due to inaccurately forecasted wind will be described. These functions ensure flying the predicted flight path. The new CDA functionalities have been implemented in DLR's 4D capable advanced flight management system (AFMS) and intensively tested during various flight trials in the A330 Full Flight Simulator in Berlin and with DLR's test aircraft ATTAS, a VFW 614 twin engine jet transport aircraft modified for research purposes. The results of these trials will be presented and discussed.

*This article describes initial results of an Airbus research project, aimed at the development of efficient operational noise abatement approach procedures that can be operated with in-service aircraft, either manually or in FMS managed mode. The project is focused on the Continuous Descent Approach (CDA) procedure, defined as an approach procedure without recourse to level segments below a height of typically 7000ft. After a first qualitative review, a noise and performance analysis has been performed for a number of CDA variants for Airbus A320 and A340 aircraft. The FMS approach procedure, including a level-off deceleration segment prior to glide slope intercept, was selected as baseline procedure. This procedure provides optimized performance and noise relief relative to conventional approach procedures. The first CDA analysis revealed a maximum descent angle, allowing convenient deceleration in clean configuration, of 2° and modest associated noise reduction relative to the baseline procedure. An optimized version of this CDA procedure, featuring a modified configuration change schedule, led to 5dBA noise relief relative to the baseline procedure. On an experimental basis, also CDA procedures with steep final glide slope angles were evaluated. Following these assessments, a series of A340 flight simulator tests was undertaken to evaluate in-flight characteristics and pilot perception for the selected CDA procedures. Simulations executed in manual mode enabled identification of acceptable descent flight path angles. Simulations of the same procedures in FMS managed mode provided further insight in operational feasibility and pilot workload issues.

An aeroacoustic study of a 26%-scale landing gear model was conducted in the NASA Ames 7- by 10-Foot Wind Tunnel using a phased microphone array. The incorporation of complex parts via stereo lithography produced a model that can mimic full-scale details down to 3 mm. These details include the contours, brake cylinders, bolt holes, and wheel hubs that appear on the real landing gear. Major noise sources were identified and ranked. From the sideline view, the noise levels of the cable harness and torque link were each at least 8 dB above that of a clean configuration. Sources from the more ambiguous fly-over view, such as the front axle, center axle and rear axle regions, were 11 dB above the clean configuration for frequencies below 2000 Hz full-scale. This increment in noise likely included other sources situated behind the truck. Referenced to the clean configuration, the braces and links contributed as much as 8 dB. Tests with a fully sealed fairing on the landing gear suggest, through careful design of major components, a noise reduction of up to 15 dB can be achieved although 2 to 6 dB of noise reduction is probably a more realistic goal. NOMENCLATURE

Calculated results of wake roll-up from an isolated transport-type wing at a low subsonic speed by a thin-layer Navier-Stokes method are presented. The computational results indicate that the vortex roll-up and its decay are faster for the flap-deployed case than for the clean configuration as indicated in experimental data. The tip vortex is relatively strong for the clean configuration than for the flapped case. The computed location of the trailing vortex for the clean configuration is near the mid-semispan where flow separation occurs, more inboard than that behind a rectangular wing. On the other hand, the trailing vortex for a flapped configuration at the same or higher lift coefficient as the clean configuration tends to be more outboard, depending on the spanwise location of the partial span flap. For a flap span with the outer edge near the mid-semispan, the outboard flap vortex merges with the trailing vortex within half span length. On the other hand, if the location of the flap outer edge is more inboard, the outboard flap vortex is still distinct even at 7.5 span lengths downstream. The condition for merging appears to be the continuity of wake vorticity distribution near the trailing edge.

A Panavia Tornado aircraft was flown to investigate airframe and jet noise in high-speed flight. The lo- cations of the noise sources were examined with a line array consisting of 29 microphones. The flights were performed in an altitude of about 35 m above ground. The influences of various parameters were studied by flying with three different airspeeds (220, 250, 275 m/s) in unaccelerated flight and three dif- ferent engine power settings (flight idle, normal, max dry) with identical airspeeds (250 m/s) over the mea- suring position. The influence of external stores on airframe noise was studied by flying with stores first and repeating all tests with an operationally clean air- craft. Two surprising results were found, (i) airframe noise is louder with the aircraft in the operationally clean configuration, and (ii) a strong noise source ex- ists close to the nozzle exit plane that dominates the noise radiation toward the side and into the forward arc.

Source distributions of airframe noise of a transport aircraft have been measured in the anechoic test section of the German-Dutch Wind Tunnel (DNW) by means of a highly directional microphone. This device - an elliptical acoustic mirrormicrophone system - had been designed for model tests in smaller wind tunnels but could be adapted to the new task. During most of the tests the model was in landing configuration with leading edge slats, flaps and landing gear extended. Some measurements of the sound source distributions were also made, for comparison, with the model in clean configuration (slats, flaps and landing gear retracted). Within the measured source distributions, dominant aerodynamic noise sources could be clearly identified. These are mainly the landing gear and the slats and flaps in extended position. The tests have demonstrated that the acoustic mirror telescope is a practical and reliable device for investigations of airframe noise sources by wind tunnel tests in the DNW.

Airframe noise has been measured on a 0.015 scale model of an advanced supersonic transport concept (AST100) in an anechoic flow facility. The model was equipped with leading- and trailing-edge flaps, nose and main landing gears, and engine nacelles. Each of these components was deployed, individually and collectively, to determine their contribution to the noise field. Results are presented which show that in the clean configuration the aircraft displays a symmetric dipole directivity, whereas in the more complex landing-approach configuration the directivity peaks in the forward quadrant. It was found that the landing-approach noise was due chiefly to the landing gear, the trailing-edge flaps, and the aeroacoustic interaction between the two.

Concepts to save fuel while preserving airport capacity by combining time based metering with profile descent procedures were developed. A computer algorithm is developed to provide the flight crew with the information needed to fly from an entry fix to a metering fix and arrive there at a predetermined time, altitude, and airspeed. The flight from the metering fix to an aim point near the airport was calculated. The flight path is divided into several descent and deceleration segments. Descents are performed at constant Mach numbers or calibrated airspeed, whereas decelerations occur at constant altitude. The time and distance associated with each segment are calculated from point mass equations of motion for a clean configuration with idle thrust. Wind and nonstandard atmospheric properties have a large effect on the flight path. It is found that uncertainty in the descent Mach number has a large effect on the predicted flight time. Of the possible combinations of Mach number and calibrated airspeed for a descent, only small changes were observed in the fuel consumed.

Airframe noise has been measured on a .015 scale model of an advanced supersonic transport concept (AST-100) in an anechoic flow facility. The model was equipped with leading- and trailing-edge flaps, nose and main landing gears, and engine nacelles. Each of these components was deployed, individually and collectively, to determine their contribution to the noise field. Results are presented which show that in the clean configuration the aircraft displays a symmetric dipole directivity, whereas in the more complex landing-approach configuration the directivity peaks in the forward quadrant. It was found that the landing-approach noise was due chiefly to the landing gear, the trailing edge flaps, and the aeroacoustic interaction between the two